•'[ij• FL •AD-A218 434

USAAVSCOM TR 89-D-22A

SAIRCRAFT "• UVSARM

SYSTEMS COMMAND

VOUE aft ESG RIEI

FED 2-••,4.3 ! SURVIVAL

AND CHECKLISTS

SIMULA INC. 4 100 dC d oIG SOUH ft~ 2ý I~ 5 r- ,1itst STREET

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DECEMBER 1989 --

FINAL REPORT

Approved for publiu rdlease; distribution is unlimited.

Prepared for

AVIATION APPLIED TECHNOLOGY DIRECTORATE

US ARMY AVIATION RESEARCH AND TECHNOLOGY ACTIVITY (AVSCOM)

FORT EUSTIS, VA 23604-5577

90 02 2 111
AVIATION APPLIED TECHNOLOGY DIRECTORATE POSITION STATEMENT

This revised -edition of the Aircraft Crash Survival Design Guide (ACSDG)

was prepared to assist those design engineers responsible for the incorporation of crashworthiness into the design of helicopters, light fixed-wing

aircraft, and tilt rotor aircraft. Also, this guide may be used in the

evaluation of the level of crashworthiness design available in the various

types of aircraft.

This report documents the components and principles of crashworthiness

and suggests specific design criteria. In general, a systems approach

is presented fur providing a reasonable level of aircrew and aircraft protection in a crash, which is considered the preferred approach. The original

Crash Survival Design Guide was published in 1967 as USAAVLABS TR 67--22

and subsequent revisions published as USAAVLABS TR 70-22, USAAMRDL TR 71-22,

and USARTL-TR-79-22A thru E. This edition consists of a consolidation

of up-t date design criteria, concepts, and analytical techniques developed

through research programs sponsored by this Directorate and others over

the past 27 years.

This document has been coordinated with other Government agencies and helicopter airframe manufacturers active in aircraft crashworthiness research

and development, and is considered to offer sound design criteria and approaches

to design for crashworthiness.

The technical monitors for this program were Messrs. LeRoy Burrows, Harold

Holland, and Kent Smith of the Safety and Survivability Technical Area,

Aeronautical Systems Division, Aviation Applied Technology Directorate.

NOTE: All previous editions of the Aircraft Crash Survival Design Guide

are obsolete and should be destroyed.

DISCLAIMERS

The findin9 in this teport are not to be conscrued a an official Department of the Army position unless so

designated by other authorized docaments.

When Government drayitigi, Specificatiorni, or other data ire used for any purpose other then in connection with %

definitely related Government procurement operation, the Unod Stetes Government thereby incurs no retpontibility

nor any obligaion whatsoe'ir; and the fact that the Government may have formulated, furnished, Or in any way

supplied the said drawings. specifications. o, ouner date is not to be regarded b'., implication or otherwise as in any

manner licensing tWe holder or any other person or corporation, or conveying any rights or permission, to manufacture, i.t, or Slll any patented invention that may in any way be related thereto.

-.red, namae cited in thic report do not constitute an official endorsement or approval of the use of such

con•n•rcial hardware or Software.

DISPOSITION INSTRUCTIONS

Destrov thit report by any method which precludes reco'structbon of the docu,, nt. ýýo not return it to the

originator.
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Pubic raporlt'g I~jriren o 'cr ¶tt (oIjcri,Ofl Oflf -rif rmtrofl q e$I S ate to averuaer hour Der retpon~e ~lang~ f e vm r rei g nf~truCtion,., searchtim ei,,tng data sOurceg atherirg and r'alntaining the data needed. afla -Imptet~ng and re~rew-nq the eallerflon Ot ýntnrnation Se'd comrit-rten reg ardintg 7his burden estimate or Ztny other tillC of thqt collection ot Informartion. trcluding 5uggestions for reduý,nq gl ottOroen. to ~V~dsh-,ttof oealouar-1ers Servlce,. Direcio'rax for inforniatorn O~eretlon and llemirts. 1215 jeflcrsr,

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December 1989 Final FROM 9/86 TO 8/89

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Aircraft Crash Survival Design Guide DAAJ02-86-C-0028

Volume I - Design Criteria and Checklsts

6. AUTHOR(S)

Richard E. Zimmermiann, No-man A. Merritt

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) S. PERFORMING ORGANIZATION

REPORT NUMBER

41 Simiula Inc,

Phoenix, Arizona 85044-5299

*9. SPONSORING/ MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. - NSORING /MONITORING

AUIENCY REPORT NUMBER

Aviation Applied Technology Directorate

U.S. Army Aviation Research & Technology Activity (AVSCOM) USAAVSCOM IR 89-D-22A

Fort Eustis, VIA 23604-5577

11. SUPPLEMENTARY NOTES

Volume I of five-volumEo report

12a. DISTRIBUTION/iAVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Approved for public release; distribut~on unlimite

( 13. ABSTRACT (Miximum 200 words)

ý- Thi s f ive-vol ure publ icati(n has been compi led to ass'i design engineers in unders;taivdinrj the design

considerations associated with !.e development of crash resistant U..Am icat olction of

avalabe ifomaton nd at pertinent to aircraft cr h eistn_ i rsneaogwt ugse

design conditions anid criteria. The five volumes of the iraf rash Sria DeinGdecover the

following topics: Volurie I - Design Criteria and Checklists; Volume 11 - Aircraft Design Crash Impact

Conditions and Human Tolerance ' Volume III - Aircraft Structural Crash Resistance; Volume 17 - Aircraft

Seats, Restraints, Litters and Cockpit'/Cabin Delethalization; and Volume V - Air(-raft Postcrash Survival.

* This volume (Volume 1) contains concise criteria drawn front Volumes II throu~h V, supplemented by

checklists intended to assist designers in implementing the criteria.

Ks 14. SUBJECT TERMS. UMBER OF PAGES

-,Arcr aft Dcsi gn buid Hfumtani Tolcran(;e fRest-aitit. Sy~tcursj'\fest Uuiminies&, )1'e2O4 Ant'Ujjtr y >PosI cra:sh sUrkvivfl' CLm I L"Ce Set 16 RIE CD ,.ra sh R s I P)ire Post!- LYSr ilu 1_F.lSy t ems Li ttcLIS .PIECD

'r tuc t i vc' Padid i i .' ~nt 7-1( u ra I C ra 11,( sics, S ILan' i, ' )______ ________

17. SEZURTY CVLASSIFICATION 18. SECURITY CLASSIFICA1IO14' 19. TSECURITY CLASSIFICATION 20. LIMITATION Or ABSTRACT

OF REPORT OF THIS PAGE OF ABSJTRACT

Ut-lLtASSIf 1-D UNICLASSIII)J1 111I ILD -

NSN 7540-011-260 5-500 Star~dord Fort', 298 (Rev 'O

IV, t ,dANiSI 'ril 139-IS
PREFACE

This report was prepared for the Safety and Survivability Technical Area of

the Aviation Applied lechnology Directorate, U.S. Army Aviation Research and

Technology Activity (AVSCOM), Fort Eustis, Virginia, by Simula inc. under Contract DAAJ02-86-C-0028, initiated in September 1986. This guide is a revision of USARTL Technical Report 79-22, Aircraft Crash Survival Design Guide,

published in 1980.

A ma.jor portion of the data contained herein was taken from U.S. Armysponsored research in aircraft crash resistance conducted from 1960 to 1987.

a Acknowledgment is extended to the U.S. Air Force, the Federal Aviation Administration, NASA, and the U.S. N~vy for their research in crash survival.

Appreciation is extended to the following organizations for provid;ng accident case histories leading to the establishment of the impact co,jitiojis in

aircraft accidents:

a U.S. Army Safety Center, Fort Rucker, Alabama.

I U.S. Naval Safety Center, Norfolk, Virginia.

* U.S. Air Force Inspection and Safety Center, Norton Air Force Base,

California.

Information was also provided by the Civil Aeronautics Board, which is no

longer in existence.

Additional credit is due the many authors, individual companies, and

organizations listed in the bibliographies for their contributions to the

field. The contributions of the following authors to previous editions of

the Aircraft Crash Survival Design Guide are most noteworthy:

D. F. Carroll, R. L. Cook, S. P. Desjardins, J. K. Drurnmond, J. L. Haley,

Jr., A. D. Harper, H. G. C. Henneberger, N. B. Johnson, G. Kourouklis, Dr. 01r. D H. Laananen, r.,r naP. . A, r . I \ r, L...... k r, 1uu . ,,. Re.. d, M. VI 3. JV ' Rely RU e IIl 'y,

S. Ht. Robertson, J. Shefrin, L. M. Shaw, G. T. Singley, III, A. E.

Tanner, Dr. J. W. Turnbow, and L. W. 1. Weinberg.

This volume was prepared ly Richard E. Zimmermann and Norman A. Merritt of

Simula Inc. Technical review and comments were provided by S. P. Uesjai'dins

of Simula Inc.

Volume I is a compilation of criteria and checkli sts for the design of crashresistant military aircraft. The criteria have been assembled in this one

volume for the conveni ence of those involved in the design or evaluation of

the overall ai rcraft and for use as a concise criteria reference. Add i tional

backlground information is provided in Volumes II through V.

1he design criteri a contLaaiined in this volume are the resul t of studies made

of crashes and experience gained during the design and manufacture of military aircraft.

iii
TABLE OF CONTENTS

Pagie

PREFACE .................... ............................... .... iii

LIST OF ILLUSTRATIONS ............ ............................ x

LIST OF TABLES ..................... ............................ xiii

INTRODUCTION ................ ............................... I

1. BACKGROUND DISCUSSION .................. ...................... 4

2. DEFINITIONS ...................... ........................... 9

- 2.1 AIRCRAFT COORDINATE SYSTEMS AND ATTITUDE PARAMETERS ......... 9

2.2 ACCELERATION-RELATED TERMS ....... ................. .... 11

2.3 VELOCITY-RELATED TERMS .............. ................... 11

2.4 FORCE TERMS ............. ......................... .... 13

2.5 DYNAMICS TERMS ............................................. 14

2.6 CRASH SURVIVABILITY TERMS ........ .................. .... 15

2.7 OCCUPANT-RELATED TERMS .............. ................... 15

2.8 SEATING GEOMET.Y 1......................18

2.9 STRUCTURAL TERMS ............ ...................... .... 20

2.10 FUEL, OIL, AND HYDRAULIC SYSTEM TERMS .............. ... 21

2.11 IGNITION SOURCE CONTROL TERMS ...... ................ .... 22

2.12 INTERIOR MATERIALS SELECTION TERMS ......... ............. 22

2.13 DITCHING AND EMERGENCY ESCAPE TERMS ........ ............. 23

3. AIRCRAFT DESIGN CRASH IMPACT CONDITIONS AND HUMAN TOLERANCE . . . 25

3.1 INTRODUCTION .............. ........................ .... 25

3.2 DESIGN CONDITIONS FOR IMPACT ..... ................ .... 25

3.2.1 Application ............................. .... 25

3.2.2 Deceleration Pulse Smape ...... ............... 25

3.2.3 Impacted Surface ...... ................... ..... 25

3.2.4 Velocity Change .............. ................... 25

3.3 HUMAN TOLERANCE TO IMPACT ............. .................. 26

3.3.1 General ........... ....................... .... 26

3.3.2 Whole-Body Tolerance ...... ................. .... 27

3.3.3 Head Injury Criteria ...... ................. .... 27

3.3.4 Spinal Injury Criteria ...... ................ .... 31

3.3.5 Leg Injury Criteria ....... ................. .... 32

3.3.6 Tolerance of Other Body Parts ............... ... 32

3.4 HUMAN BODY I)IMENSIONS AND MASS DISTRIBUTIONS .......... ... 32

3.4.1 General ........... ....................... .... 32

3.4.2 Anthropometry ......... .................... .... 32

3.4.3 lnerl.ial Properties ....... ................. .... 35

3.5 CRASH IEST DUMMIES ......... ..................... ... 35

V
TABLE OF CONTENTS (CONTD)

4. AIRFRAME STRUCTURAL CRASH RESISTANCE ......... ............... 37

4.1 INTRODUCTION .............. ........................ 37

4.2 AIRFRAME CRASH RESISTANCE ........... ................. 41

4.2.1 Longitudinal Impact ..... ................. .... 41

4.2.2 Vertical Impact ...... ................... .... 43

4.2.3 Lateral Impact ...... ................... .... 43

4.2.4 Rollover Impact ...... ................... .... 44

4.2.5 Wings and Empennage ..... ................. .... 45

4.2.6 Engine/Transmission Mounts .... ............. .... 45

4.2.7 Landing Gear ........................ 46

4.3 ANCILLARY EQUIPMENT RETENTION .... ............... ..... 47

4.3.1 Strength .......... ...................... .... 47

4,3.2 Emergency and Survival Equipment Stowage Location. 4-

4.3.3 Release of Emergency and Survival Equipment ........ 48

4.4 INTERFACE OF OCCUPANT AND CARGO RETENTION SYSTEMS

WITH AIRFRAME ............... ....................... 48

4.5 OCCUPANT RETENTION ............ ..................... 49

4.6 CARGO RETENTION .......... ...................... .... 49

4.7 SPECIAL TILT-ROTOR CONSIDERATIONS .... ............. .... 50

4.8 TESTING ... .......................... 53

4.8.1 Aircraft System Testing ......... ............... 53

4.8.2 Landing Gear Crash Testing .... ............. .... 53

4.8.3 Cargo Restraint ...... ................... .... 54

4.8.4 Seat and Restraint System ....... .............. 54

4.8.5 Fuel System ...... ..................... ..... 54

4.8.6 Ancillary Equipment Retention ............... .... 54

4.9 DESIGN CHECKLISTS ................................. ...... 54

4.9.1 Landing Gear Design Checklist ............... .... 54

4.9.2 Airframe Design Checklist ......... .............. 55

5. AjRCRAFT SEATS, RESTRAINTS, LITTERS, AND COCKPIT/CABIN

DELETHALiZATiON ............................................. .

5.1 INTRODUCTION ..... ........................ 58

5.2 PRIMARY DESIGN CONSIDERATIONS .... ............... ..... 58

5.2.1 General ... ....................... 58

5.2.2 Design Conditions ........... .................. 58

5.2.3 Structural Distortion ..... ................ .... 58

5.3 DESIGN PRINCIPLES FOR SEAIS AND LITTERS ............. .... 59

5.3.1 Seating System Orientation .... ............. .... 59

5.3.2 Litter Orientation ........... ................. 60

5.3.3 Materials ....... ...................... .... 60

vi
TABLE OF CONTENTS (CONTD)

Page

5.4 STRUCTURAL CONNECTIONS ...... ................... .... 60

5.4.1 Bolted Connections ..... ................. ..... 60

5.4.2 Riveted Connections ..... ................. .... 61

5.4.3 Welded Connections ..... ................. ..... 61

5.4.4 Seat Attachment ...... ................... .... 61

5.4.5 Joint Deformation .......... .................. 62

5.4.6 ?'aterial .......... ...................... .... 62

5.4.7 Restraint System Anchorage .... ............. .... 66 d 5.5 ENERGY-ABSORBING DEVICES .......... .................. 66

5.6 SEAT CUSHIONS .......... ....................... ..... 68

5.6.1 General ....... ....................... ..... 68

5.6.2 Requirements ............. .................... 69

5.6.3 Energy-Absorbing Cushions .................. .... 69

5.6.4 Net-Type Cushions .......... .................. 69

5.6.5 Seat Back Cushions ........... ................. 69

5.6.6 Headrests .......................... 69

5.7 DESIGN PRINCIPLES FOR PERSONNEL RESTRAINT SYSTEMS ..... 70

5.7.1 General ....... ....................... ..... 70

5.7.2 Types of Systems ..... .................. ... 70

5.7.3 General Design Criteria ................... ..... 74

5.7.4 Webbing and Attachments ................. 82

5.7.5 Inertia Reels, Control, and Installation .... ...... 86

5.8 SEAT STRENGTH AND DEFORMATION DESIGN REQUIREMENTS ..... 86

5.8.1 Recommended Occupant Weights for Seat Design . . .. 86

5.8.2 Strength and Deformation .......... .............. 88

5.8.3 Other Seats ...................................... 92

5.9 PERSONNEL RESTRAINT HARNESS TESTING .... ............ .... 93

5.10 STRUCTURAL SYSTEM TEST REQUIREMENTS .... ............ .... 93

5.10.1 Static Test Requirements .................. .... 94

5.10.2 Dynamic Test Requirements .... ............. .... 99

5.10.3 Data Acquisition and Reduction ...... ........... 102

5.10.4 Seat Component Attachment ............. 104

5.11 RETROFIT FOR SEATING SYSTEMS ..... ................ .... 104

5.11.1 General .......... ...................... .... 104

5.11.2 Forward-Load-Limiting Seats .... ............... 105

5.12 LITTER STRENGTH AND DEFORMATION REQUIREMENTS ..... ........ 106

5.12.1 General .......... ...................... .... 106

5.12.2 Recommended Occupant Weights for Litter Design. . 106

5.12.3 Vertical Loads ......................... ..... 106

5.12.4 Lateral and Longitudinal Loads ...... ........... 108

5.12.5 Litter Restraint Harness Testing ............. .... 109

5.12.6 Litter System Test Requirements ....... ..... .... 109

Accesion F-r

NTIS CY:J,,& d

CD'C 1E.L4I

D, .L 

vii
TABLE OF CONTENTS (CONTD)

5.13 DELETHALIZATION OF COCKPIT AND CABIN INTERIORS ........... 110

5.13.1 General .............................. 110

5.13.2 Occupant Strike Envelopes ...... ............. .11

5.13.3 Head Strike Envelope in Stroking Seats ........... 113

5.13.4 Environmental Hazards ..... ............... . 113 !.. 

5.13.5 Head Impact Hazards ...... ................ ... 115

5.13.6 Instrument Panel Structure Proximity ..... ........ 115

5.13.7 Rudder Pedal Configuration ................. ... 116

5.13.8 Controls and Control Columns ............... .... 116

5.13.9 Sighting and Visionic Systems .... ........... 117

5.13.10 Energy-Absorbing Requirements for Cockpit and

Cabin Interiors ........... .................. 118

5.14 DESIGN CHECKLISTS ..................... 125

5.14.1 General Design Chec~list .................. .... 125

5.14.2 Seat Strength and Deformation Checklist .... ...... 126

5.14.3 Seat Cushions Checklist ................... 126

5.14.4 Litter Strength and Deformation Requirements

Checklist ........................... 127

5.14.5 Restraint System Design Checklist. ........... .... 127

5.14.6 Protective Padding Checklist ................ ... 129

5.14.7 Cockpit Controls and Equipment Checklist .......... 129

6. AIRCRAFT POSTCRASH SURVIVAL ...... ................... .... 130

6.1 INTRODUCTION ......... ........................ .... 130

6.2 FUEL SYSTEM DESIGN CRITERIA ..... ................ .... 130

6.2.1 General ....... ....................... .... 130

6.2.2 Fuel Tanks ...... ....... ..................... 130

6.2.3 Fuel Lines ...... ....... ..................... 131

6.2.4 Frangible Attachments ..... ................ ... 134

6.2.5 Self-Sealing Breakaway Valves ................ ... 136

6.2.6 Fuel Drains ............ ..................... 136

6.2.7 Filler Units . . .. ... . I I.. .... . 1138

6.2.8 Fuel Boost Pumps ........... .................. 139

6.2.9 Fuel Filters and Strainers .... ............. .... 139

6.2.10 Fuel Valves ....... ..................... .... 140

6.2.11 Fuel Quantity Indicators .......... .............. 140

6.2.12 Vents ......... ........................ .... 140

6.3 OIL AND HYDRAULIC SYSTEM DESIGN CRITERIA ............. .... 141

6.3o1 General ....... ....................... .... 141

6.3.2 Oil and Hydraulic Fluid Reservoirs ............ ... 141

6.3.3 Oil and Hydraulic Lines ....... ............... 142

6.3.4 Oil and Hydraulic System Components ........... .... 143

6.3.5 Oil Coolers ............................ 143

6.4 IGNITION SOURCE CONTROL CRITERIA ....... .............. 143

6.4.1 Electrical Systems ...... ................. .... 143

6.4.2 Shielding ......... ...................... ... 145

viii
TABLE OF CONTENTS (CONTD)

Page

6.5 INIERIOR MATERIALS SELECTION CRITERIA .............. .... 145

6.5.1 General ....... ....................... .... 145

6.5.2 FAR 25.853 Flammability Requirements ........... ... 145

6.5.3 Smoke and Toxic Gas Test Criteria ............. .... 146

6.6 DITCHING CRITERIA ........ ..................... ... 147

6.6.1 General ....... ....................... .... 147

6.6.2 Emergency Exits ...... ................... ... 147

6.6.3 Underwater Emergency Lighting ................ ... 147

6.6.4 Helicopter Flotation Systems .... ............ ... 147

6.6.5 Ditching Equipment ..... ................. .... 148

6.7 EMERGENCY ESCAPE AND RESCUE DESIGN CRITERIA ........... ... 148

6.7.1 Emergency Exits ...... ................... ... 148

6.7.2 Emergency Lighting .......... ................. 151

6.7.3 Emergency Exit Markings ....... ............... 152

6.7.4 Crew Chief Stations ..... ................. .... 153

6.7.5 Alarm Systems ...... .................... .... 153

6.8 CRASH LOCATOR BEACON DESIGN CRITERIA ................ ... 153

6.8.1 General ....... ....................... .... 153

6.8.2 Crash Sensors ...... .................... .... 153

6.8.3 Transmitters . . . . . . . . . . .. . 1.. !54

6.8.4 Antennas ....................... 155

6.8.5 Power Supplies ...... ................... ..... 55

6.9 DESIGN CHECKLISTS .... ................................. 155

6.9.1 Fuel System Design Checklist .... ............ ... 155

6.9.2 Oil and Hydraulic System Design Checklist .......... 161

6.9.3 Ignition Source Control Checklist ............ ... 163

6.9.4 Interior Materials Selection Checklist .... ....... 164

6.9.5 Ditching Provisions Checklist ................ ... 165

6.9.6 Emergency Escape Design Checklist ............. .... 165

6.9.7 Crash Locator Beacon Checklist ...... ........... 168

REFERENCES ....................... ...... .............................. 170

BIBLIOGRAPHY ............ ............................. ... 177

ix
LIST OF ILLUSTRATIONS

Figure Page

1 Energy management system ...... .................. . ... 1

2 Procedure for evaluation of structural designs with respect

to crash resistance ....... ..... ..................... 7

3 Aircraft coordinates and attitude directions ..... ........ 9

4 Typical aircraft floor acceleration pulse ............. .... 12

5 Terminology for directions of forces on the body .... ...... 16

6 Seating geometry .......... ...................... ... 18

7 Duration and magnitude of spineward acceleration endured

by various subjects ....... ... ..................... 28

8 Initial rate of change of spineward acceleration endured

by various subjects ....... ... ..................... 29

9 Duration and magnitude of headward acceleration endured

by various subjects ................. .... 30

10 Initial rate of change of headward acceleration endured

by various subjects ....... ... ..................... 31

11 Conventional seated anthropometric dimensions ..... ........ 33

12 Structural layout for occupant protection during crash

impact conditions ......... ...................... ... 40

13 Low-angle impact design conditions (simulated approach

with antitorque loss under poor visibility) ........... .... 42

14 Nose section design conditions ..... ............... ... 42

15 Rollover, roof impact design condition ...... ........... 44

16 Rollover, side impact design condition ...... ........... 45

17 Load-displacement requirements for energy-absorbing cargo

restraint systems (forward loading of rotary-and fixed-wing

aircraft) ........... .......................... ... 51

]8 Cargo lateral load-displacement requirements ..... ........ 52

19 Wing/rotor/pylon crash failure modes .... ............ ... 53

x
LIST OF ILLUSTRATIUNS (CONTD)

Figure Lo

20 Floor or bulkhead warpage requirement for static loading

of seat .......... ..... ........................... 63

21 Bulkhead in-plane warpage ...... .................. ... 64

22 Pin joint releases oriented to allow rotation around an

aircraft roll axis ....... ..................... ... 65 4I

23 Basic aircrew restraint system ......... ............... 71

24 Aircrew restraint system, including reflected shoulder

straps ........... ........................... .... 72

25 Aircraft troop/passenger restraint systems ..... ......... 73

26 Gunner restraint system .......... ................... 74

27 Inflatable body and head restraint .... ............. 75

28 Buckle fitting at.,.•hmvnt and motion angles ..... ......... 76

29 Lap belt anchorage _2ometry ..... ................. .... 78

30 Shoulder harness anchorage geometry ....... ............. 80

31 Stitch pattern and cord size ......... ................ 84

32 Wrap radius for webbing joints ..... ............... .... 85

33 Webbing fold at metal hardware attachment ..... .......... 85

34 Seat forward load and deflection requirements for all

types of Army aircraft (forward design pulse) ..... ........ 89

35 Lateral seat load and deformation requirements for all

types of Army aircraft ........... ................... 93

36 Static load application point and critical body block

pelvis geometry ........ ... ....................... 96

37 Dynamic test requirements for qualification ..... ......... 98

38 Example of input pulse for seats having less than 12 in.

of stroke ........ ....... .......................... 101

39 Graphic approximation example ....... ................ 104

xi
LIST OF ILLUSTRATIONS (CONTD)

Figure PacwP

40 CH-53 crew seat ....... ....................... ... 105

41 Litter downward load and deflection requirements .... ...... 107

42 Litter forward or lateral loaa and deflection requirements

for all types of Army aircraft ..... ............... ... 108

43 Full-restraint extremity strike envelope - side view .... 111 k

44 Full-restraint extremity strike envelope - to, view. ....... 112

45 Full-restraint extremity strike envelope - front view .... 112

46 Lap belt-only extremity strike envelope - side view .... 113

47 Lap belt-only extremity strike envelope - top view ........ 114

48 Lap belt-only extremity strike envelope - front view .... 114

49 Wayne State tolerance curve for the human brain in

forehead impacts against plane. unyi•!ding surfaces. . . . . 116

50 Antitorque, or rudder, pedal geometry to prevent

entrapment of feet ........ ..................... ... 117

51 Crushable eyepiece concept ........... .................. 119

52 Recommended stress-strain properties for padding material

for head contact with cushion thickness of at least

1.5 in .................. .......................... 122

53 Effect of density on stress-strain curves for polyurethane- fdig,q•u •

54 Hose assembly test modes ........................ .. ... 132

55 Sample frangible attachment separation load calculation. . 135

56 Typical method of breakaway load calculation for fueltank-to-line breakaway valve ..... ................ ... 137

57 Static tensile and shear test modes for self-sealing

breakaway valves .......... ...................... .. 138

xii
LIST OF TABLES

Table Page

1 Crash-resistance criteria for the preliminary design

process ....... ..... ... ........................... 8

2 Crash impact design conditions with landing gear extended,

MIL-STD-1290 ........ ........................ .. 26

3 Summary of anthropometric data for U.S. Army male

aviators ...... ..... .......................... ... 34

4 Summary of anthropometric data for male soldiers .... ...... 34

5 Performance requirements under crash impact conditions

per Table 2 ......... ......................... ... 38

6 Other structural performance requirements ..... .......... 39

7 Cargo restraint loads and displacement requirements. ..... 50

8 Aircraft cargo categories ...... .................. ... 52

9 Occupant restraint hdrness requirements ..... ........... 82

10 Minimum webbing width requirements ...... ............. 83

11 Typical aviator weights ...... ................... .... 87

12 Troop and gunner weights ...... .................. ... 87

13 Seat design and static test requirements ... .......... 95

14 Data channels ....... ....... ........................ 102

15 Litter system static test requirements .... ........... ... 109

16 Summary of ASTM test methods and specifications for,

flexible cellular plastics ...... ... ................. 121

17 Required minimum individual loads for standard hose and

hose-end fitting combinations ..... ................ ... 133

xiii
TNTRODUCTION

For many years, emphasis in military aircraft accident investigation was placed

on determining the cause of the accident. Very little effort was expended on

the crash survival aspects of aviation safety. However, it became apparent

through detailed studies of accident investigation reports that significant

improvements in crash survival could be made if consideration were given in the

initial aircraft design to the following factors that influence survivability:

1. Crash Resistance of Aircraft Structure - The ability of the aircraft

structure to maintain living space for occupants throughout a crash.

2. Tiedown Strength - The strength of the linkage preventing occupant,

cargo, ur equipment from breaking free and becoming missiles during

a crash sequence.

3. Occupant Acceleration During Crash Impact - The intensity and

duration of accelerations experienced by occupants (with tiedown

assumed intact) during a crash.

4. Occupant Crash Impact Hazards - Barriers, projections, and loose

equipment in the immediate vicinity of the occupant that may cause

contact injuries.

5. Postcrash Hazards - The threat to occupant survival posed by fire,

drowning, entrapment, exposure, etc., following the impact sequence.

Early in 1960, the U.S. Army Transportation Research Command* initiated a

long-range program to study all aspects of aircraft safety and survivabilty.

Through a series of contracts with the Aviati*'n Safety Engineering and Research (AvSER) Division of the Flight Safety Foundation, Inc., the problems

associated with occupant survival in aircraft crashes were studied to determine specific relationships among crash forces, structural failures, crash

fires, and injuries. A series of reports covering this effort was prepared

and distributed by the U.S. Army, beginning in 1960. In October 1965, a

special project initiated by the U.S. Army cnnsnlidated the design criteria

presented in these reports into one technical document suitable for use as a

designer's guide by military aircraft design engineers. The document was to

be a summary of the cu-rent state c.f the art in crash survival design. The

Crash Survival Design Guide, TR 67-22, published in 1967, realized this goal.

Since its initial publication, the Design Guide has been revised and expanded

four times to incorporate the rpsults of continuing research in crash resistance technology. The third edition, publish.ed in 1971. wa,- the basis for

the criteria contained in the original version of the Army's military standard MLL-STD-1290, "Light Fixed-- and Rotary-Wing Aircraft Crash Resistance"

(Reference 1). The fourth edition, pubiished in 1980, entitled Aircraft

Crash Survival DesignGuide, expanded the dociment to five volumes, which11

have been updated by the current edition to include information and changes

*Now the Aviation Applied Technology Directorate, U.S. Army Aviation Research

and Technology Activity, U.S. Army Aviation Systems Command (AVSCOM).
developed from 1980 to 1987. This current edition, the fifth, contains the

most comprehensive treatment of all aspects of aircraft crash survival now

documented. It can be used as a general text to establish a basic understanding of crash impact conditions and the techniques that can be employed

to improve chances for survival. It also contains design criteria and checklists on many aspects of crash survival and thus can be used as a source of

design requirements.

It should be emphasized that the Design Guide is to be used as a guide, not as

a specification. System specifications should reference applicable crashresistant design specifications, such as MIL-STD-1290, MIL-S-58095, and

MIL-S-85510, or should include specific criteria selected from the Design

Guide or other sources.

The current edition of the Aircraft Crash Survival Design Guide is also

published in five volumes. Volume titles and general subjects included in

each volume are as follows:

Volume I - Desiqn Criteria and Checklists

Pertinent criteria extracted from Volumes II through V, presented in the

same order in which they appear in those volumes.

Volume II - Aircraft Design Crash Impact Conditions and Human Tolerance

Crash impact conditions, human tolerance to impact, military

anthropometric data, occupant environment, test dummies, accident

information retrieval.

Volume III - Aircraft Structural Crash Resistance

Crash load estimation, structural response, fuselage and landing gear

requirements, rotor requirements, ancillary equipment, cargo restraints,

structural modeling.

Volume IV - Aircraft S?ats, Restraints, Litters, and CockpitLCabin 'D eI ..... hl .. . . . . . . DILfldI I di UII

Operational and crash impact conditions, energy absorption, seat design,

litter requirements, restraint system design, occupant/restraint system/

seat modeling, delethalization of cockpit and cabin interiors.

Volume V - Aircraft Postcrash Survival

Postcrash fire, ditching, emergency escape, crash locator beacons.

In this volume (Volume 1), Chapter I introduces and explains the intended use

of the material contained herein. Chapter 2 contains definitions of terms

used it) the Design Guide. Chapters 3, 4, 5, and 6 contain the criteria and

checklists extracted from Volumes II, II1, IV, and V, respectively. The

reader of this volume is strongly urged to familiarize himself with the

material in the other volumes, at least in his particular area of responsibility (for example, seats and restraints or fuel systems), in order to

appreciate more fully the limitations of the criteria.

2
The criteria are supplemented by checklists that are intended for use by aircraft designers in the original design stages and in the design review. These

checklists should help the designer apply the necessary criteria in a comprehensive and orderly manner during the development of crash-resistant designs

and provide a rapid and positive means of determining that none of the criteria have been overlooked. The responses on the checklists also should aid

the designer in determining the strengths and weaknesses of an existing or

proposed design.

After the designer has finished reviewing a system design, each item on the

applicable checklists should have a check mark in one of the spaces following

the item. Those items marked "NO" should be examined to determine the reason

for noncompliance with the design criteria. Unless the reason involves a conflicting, overriding requirement, the design should be revised to meet the

crash-resistant criteria. Those items marked "N/A" should be carefully reviewed to be sure that the item is truly not applicable to the system under

consideration.

The units of measurement shown in the Design Guide vary depending upon the

units used in the referenced sources of information, but are mostly USA

units. In some cases the corresponding metric units are shown in parentheses

following the USA units. For the convenience of the reader a conversion table

of some commonly used units follows:

USA Unit Abbr. or S•pibol Metric Equivalent Abhr. or Symbol

We ioIt

Ounce oz 28.350 grams g

Pound lb or # 0.454 kilogram ko

Capacity

(U.S. liquid)

Fluidounce fl oz 29.573 milliliters i1l

Pint pt 0.473 liter

Quart qt 0.946 liter

Gallon gal 3.705 liters

* Length

Inch in. 2.54 centimeters cm

Foot ft 30.48 centimeters cm

Yard yd 0.9144 meter m

Mile mi 1.609 kilometers km a

Area S~2

Square Inch sq in. or in. 6.492 sq centimctr.r sq cm or ciL

Square loot sq It or ft 0.093 squaron moter sq m or m2

Vo I ume

Cubic Inc.h (u in. ur in. 3 IG.3657 LUub1,• L ntinicterz cu Lio or rni3

Cubic Foot cu ft or ft3 0.028 cul.bIc meter tu m o, mn3

Forte

Pound lb 4.446? x 10r dy'us

4.441j? r"owl t ui,

3
1. BACKGROUND DISCUSSION

lhe overall objective of designing for crash resistance is to eliminate injuries and fatalities in relatively mild impacts and minimize them in all severe

but survivable mishaps. A crash-resistant aircraft will also reduce aircraft

crash impact damage. By minimizing personnel and material losses due to crash

impact, crash resistance conserve:; resources, is a positive morale factor, and

improves the effectiveness of the fleet both in peacetime and in war. Results

from analyses and research have shown that the relatively small cost in dollars and weight of including crash resistance features is a wise investment

(References 2 through 13). Consequently, new-generation Army rotary-wing aircraft are being procured to stringent, yet practical, requirements for crash

resistance.

To provide as much occupant protection as possible, a systems approach to

crash resistance must be followed. The systems approach to crash resistance

means that the landing gear, aircraft structure, and occupant seats must all

be designed to work together to absorb the aircraft kinetic energy and slow

the occupants to rest without injurious loading, as shown in Figure 1 (Reference 14). In addition, the occupants must all be restrained and a protective

structural shell maintained around the occupied areas during a crash to provide a livable volume. Weapon sights, cyclic controls, glare shields, instrument panels, armor panels, and aircraft structure must be delethalized if

they lie within the strike enveiope of the occupant. Postcrash hazards, such

as fire, entrapment, drowning, emergency egress, and rescue must also be

considered in an effective crash-resistant design.

LARGE MASSES SLOWED

D OWN BY GEAR AND

FUSELAGE

SEA -... OCCUPANT SLOWLD DOWN

FUSEI.AGE BY GLAR, FUSELAGE.

[ •]--,]AND SEAT

LANUING GEAR

FIGURE 1. ENERGY MANAGEMLNT SYSTEM. (FROM RLFERENCE 14)

4
The composition of the surface being impacted must also be considered early in

the design phase, when decisions concerning the relative energy-absorbing

roles of the landing gear, structure, and seats are made. When ri:pacting on

hard surfaces, landing gear with high energy-absorbing capacity can protect

the fuselage from major damage during low-velociy impacts and provide occupant protection during higher-velocity impacts. However, during impacts with

soft surfaces, such as water or marshy ground or uneven surfaces caused by

rocks, trees, etc., the force acting on the landing gear may not be great

enough to activate its energy-absorption function, and it will not contribute

at all to the occupant protective system. Aircraft with skids, instead of

wheels, could provide better protection in such impacts, because the greater

surface area of the skids would transmit a greater load and possibly activate

the landing gear energy absorbers. The surface conditions can also affect the

severity of the impact. Soft soil can deform and contribute energy absorption

during a vertical impact. Soft soil can also cause plowing and rapid deceleration if there is a large horizontal velocity component. A vertical impact on

water can be very severe, since the large aircraft fuselage area contacting

the water would result in high deceleration rates and various crash load

paths. A high-speed longitudinal impact into water can also cause high loads

from water plowing as water enters through lower nose transparencies.

Thus it is usually not possible to design a system that relies only on one or

two of the three energy absorption features available (landing gear, structure, and seats). Now that more helicopters use retracting landing gears,

more emphasis orn energy-absorbing fuselage understructure is required to corntrol loads and permit the energy-absorbing seats to function. Even though it.

is difficult to design for, and predict the behavior of, energy-absorbing fuselage structure, it is as important as the other components in the system.

The introduction of composite primary structure into modern aircraft presents

special problems for the designer dealing with crash rcsistance. The brittle

failure modes of most composites makes the design of energy-absorbing crush

able structures difficult, but not impossible. fortunately, other compositi,

materials are suitable for use in such fuselaye structures. A more detailed

discussion of designing with composite. is presented in Volume I11.

It would seem efficient to simply specify human toler anci, reqnirerments and an

array of vehicle crash impact conditions and then develop the helicopter as a

crash-resistant system with an efficient mixture of crash-resistant. features

for that particular helicopter. However, available structural and human tol

erance analytical techniques needed to perform, evo, uate, and validate such a

maximum design freedom approach to achieving crash resistance are riot suf iciently (:omprehensive to b,, relied upon comiiietely. l urthrmore, testing corn

plete aircraft sufficiently early in the developmnent cycle to permit evallia

tion of system concepts is riot practical. lhe systems approach dictates that

the designer consider probable crash conditiuns wherein one or more subsysteiis

do riot perform their desired functiours, for example, an irippa, . sitnation in

which the landing ye!ar does riot abisorb its share of the ilwajrt crash C:•lY

because of aircraft attitude at irriact. lhirefore, to achieve the overal I

nol , ininimum levels of (.rash proute.ction aye rc.omirwriidud for the varionus

individual subsystturrs with ihalance bu tween thu two extrermres of" (I) def ining

necessary performance on a comprrnetrt level only, aid (2) r,;quirirgn that tlre

aircraft. system bhe desi gnied onýIly fur i rrpiatt cumdlitioro . with 110o currponr:I.

design and test criteria.
Current aircraft cr'ash resistance criteria require that a new aircraft be designed as a system to meet the vehicle impact design cornditions recommended in

"dc,,mie U,; however, minimum (ri uicia arp also speci `4ed for a number of crash.-

critical c:ipnneprils. For .xýrple, ninn:mur crasr energy-absoruti)ýn requirements

for seats and restraint anyst;.s ,,d larndiv, oear ,•ip specifieu. All strength

requirements presented in this volume are based on the crash impact conditions

described in Volume IT. Testing requirements are based on ensuring compliance

with strength and deformation requirements. Crash-resistance design criteria

for U.S. Army light fixed- and rotary-wing aircraft are stated in MIL-STD-1290

(Reference 1). All pilot, copilot, observer, and student seats in these aircraft should conform to the requirements of MIL-S-58095 (Reference 15), while

passenger seats should conform to MIL-S-85510 (Reference 16).

Although much higher levels of crash resistance can be achieved during the development of completely new aircraft designs, the crash resistance or existing

aircraft can be significantly improved through retrofitting these aircraft with

crash-resistant components adhering to the design principles of this design

guide. This can even be achieved while expanding the combat effectiveness of

the aircraft. Examples of this are the successful programs to retrofit all

U.S. Army helicopters with crash-resistant self-sealing fuel systems (Reference 17) and the U.S. Navy program to retrofit the CH-46 SH-3, HH-3, and CH-53

helicopters (References 18, 19, and 20) with crash-resistant armored crewseats.

In an initial assessment, the definition of an adequate crash-resistant structure may appear to be relative!y simple. In fact, many influencing parameters

must he considered before an optimum design can be finalized. A complete systems approach (as summarized in Figure 2) should be employed to include all

influencing parameters concerned with the design, manufacture, overall performance, and economic constraints on the aircraft in meeting mission requirewents. lrade-offs among the parameters must be made in order to arrive at a

final design that most closely meets the customer's specifications. Each type

of aircraft may require a different emphasis in the parameter mix. Table I

summarizes major crash resistance criteria that should be considered during the

preliminary design phase.

r

6
4 ore

Analytical method Structure Occupant environment

to determine utput response Yes response

structure and

occupant responses < criteria • <Criteria Reduce

~-IrputI

In u Ne structure

Require- load-deformation

ments Results characteristics

Probabule Crash - Fnvironment Linicnar and Incremental

TakeroI f C|,-* utul sties Acceleration ht C

T~akeoFs Cbazaterssjc Duration

Land]"g Fuselage Volume change Cruise ettc

Gear Dclethalization _No Acceptable

oalli n ingine_ weight and Stall Froing t r space change

withU ground Tai 1 Ldad

(sentd(u Nm, Accelptable Design

C o ] tist j n w it h s ura M o m e n t I A ,t C -

-e- is.o with --- i Deflection Impact ConditionsI

Trrainu liending Deformation *l'renslation and

Terrain type Shear Failure mode rotational velocities

and slop, lluckli•n Encrgy and rate-s

---- (I,-sVnsevere -I • vij • Int) II____.........._________ = Attitude,

oTerla'n type and slope

FIGURE 2. PROCEDURE FOR EVALUATION OF STRUCTURAL DESIGNS

WITH RESPECT TO CRASH RESISTANCE.

7
TABLE 1. CRASH-RESISTANCE CRITERIA FOR THE PRELIMINARY DESIGN PROCESS

Postcrash

Crash Scenarios Primary Structure Energy ALsorption Requirements

M MIL-STD-1290 * Support of large a Landing gear a Emergency egress

defines predom- mass items

inant impact a Controlled struc- - Occupant release

conditions a Support of sys- tural collapse from seats

tems

* Single axis and * Crash-resistant - Door/exit

combination of: * Occupant support energy-absorbing openirj

and protection seats

* Shedding of large - Accessibility

- Vertical impact & Cargo contain- mass items: and illumination

ment and tiedown of exits

- Longitudinal - Engines

impact * Support of land- a Minimization of

ing gear loads - External stores postcrash tire

- Lateral impact hazards

* Space consistent - Tail boom

Post impact with occupant - Fuel containment

Rollover strike envelope (Shed items must

Pitchover not impact occu- - Oil and hydraulic

Nose plowing a Emergency exit pied areas) fluid containment

structure

a Controlled - Fuel modification

displacement of:

- Ignition source

- Transmissions control

- Rotor heads a Reduced material

flammability,

* Impacted surface smoke and

(soft ground etc.) toxicity @

8|
2. DEFINITIONS

2.1 AIRCRAFT COORDINATE SYSTEMS AND ATTITUDE PARAMETERS

0 Aircraft Coordinates

Positive directions for velocity, acceleration, and force components

and for pitch, roll, and yaw are illustrated in Figure 3. When referring to an aircraft in any flight attitude, it is standard practice to use a basic set of orthogonal axes as shown in Figure 3,

with x, y, and z referring to the longitudinal, lateral, and vertical directions, respectively.

+z

YAW

+ X ROLL X

NOTE: RIGHT-HAND RULE DOES NOT APPLY.

FIGURE 3. AIRCRAFT COORDINATES AND ATTITUDE DIRECTIONS.

However, care must be exercised when analyzing ground impact cases

where structural failure occurs, aircraft geometry changes, and

reaction loading at the ground plane takes place. In the simulation

of such impacts, it is often necessary to use more than one set of

reference axes, including the earth-fixed system shown in Figure 3

as X, Y, Z.

9
0 Attitude at Impact

The aircraft attitude, with respect to the aircraft coordinate sys--

tem, in degrees at the moment of initial impact. The attitude at

impact is stated in degrees of pitch, yaw, and roll (see Figure 3).

Aircraft pitch is the angle between its longitudinal axis and a

horizontal plane. Pitch is considered positive when the nose of the

aircraft points above the horizon and negative when it points below

the horizon. Yaw is measured between the aircraft's longitudinal

axis and the flight path. Roll is the angle between an aircraft

lateral (y) axis and the horizontal, measured in a plane normal to

the aircraft's longitudinal axis.

0 Flight Path Angle

The angle between the aircraft flight path and the horizontal at the

moment of impact.

• Terrain Angle

The angle between the impact surface and the horizontal, measured in

a vertical plane.

* Impact Angle

The angle between the flight path and the t-,riin, measured in a

vertical plane. The impact angle is the alg9, 1aic sum of the flight

path angle plus the terrain angle.

"• • •H• PTH •IMPACT ANGLE

UPHILL FLIGHT

FLIGHT PATH

IMPACT

ANGLE

-FLIGHT PATH TERRAIN

HORIZONTAL

DOWNHILL FLIGHT

10
2.2 ACCELERATION-RELATED TERMS

* Acceleration

The rate of change of velocity. An acceleration is required to

produce any velocity change, whether in magnitude or in direction.

Acceleration may produce either an increase or a decrease in velocity. There are two basic types of acceleration: linear, which

changes translational velocity, and angular (or rotational), which

changes angular (or rotational) velocity. With respect to crash impact conditions, unless otherwise specified, all acceleration values

are those at a point approximately at the center of the floor of the

fuselage or at the center of gravity of the aircraft.

* Deceleration

Acceleration in a direction to cause a decrease in velocity.

a Abrupt Accelerations

Accelerations of short duration primarily associated with crash

impacts, ejection seat shocks, capsule impacts, etc. One second is

generally accepted as the dividing point between abrupt and prolonged accelerations. In abrupt accelerations the effects on the

human body are limited to mechanical overloading (skeletal and soft

+4ssue .. . L)s . .llLIIr.. . li'li . L r1L time for functional disturbances due to fluid shifts.

* The Tern, G

The ratio of a particular acceleration (a) to the acceleration (g)

due to gravitational attraction at sea level (32.2 ftisec );

G = a/g. In accordance with common practice, this report will refer

to accelerations measured in G. To illustrate, it is customarily

understood that 5 G represents an acceleration of 5 x 32.2, or

161 ft/sec2 .

e Rate of Onset

Rate of application of G's, expressed in G's per second (rate of

change of acceleration).

4G

Rate of Onset = t (G's per second)

2.3 VElOCITY-RELATED TERMS

0 Velocity Chanqe in Major Impact (Av

The decrease in velocity of the airframe during the major impact,

expressed in feet per second. The major impact is the one in which

the highest forces arc incurred, not necessarily the initial

impact. For the acceleration pulse shown in Figure 4, the major

impact should be considered ended at time t 2. Elastic recovery in

11
P L A K - --_--__

z

0

i- AVERAGE T I II

w~ I

w I

tII

FIGURE 4. TYPICAL AIRCRAFT FLOOR AC;CELERATurN PULSE.

the structure will tend to reverse the direction of the aircraft

velocity prior to t 2 . Should the velocity actually reverse, its direction must be considered in computing the velocity change. For

example, an aircraft impacting downward with a vertical velocity

component of 30 ft/sec and rebounding with an upward comporent of

5 ft/sec should be considered to experience a velocity change

v= 30 - (-5) -35 ftsec

during the major impact. The velocity change during impact is

further explained in Section 7.1 of Volume alr.

• Longitudinal Velocity Chancie

The decrease in velocity during the major impact measured a~ong the longitudinal (roll) axis of the aircraft. The vellcity may or may

not reach zero during the major impact. For examplc, an aircraft

impacting the ground at a forward velwcity of 100 ft/sic and slowing

to 35 ft/sec would experience a longitudinal velocity change of

65 ft/sec during this impact.

12
* Vertical Velocity Change

The decrease in velocity during the major impact measured along the

vertical (yaw) axis of an aircraft. The vertical velocity generally reaches zero during the major impact and may reverse if rebound

occurs.

0 Lateral Velocity Change

The decrease in velocity during the major impact measured along the

lateral (pitch) axis of the aircraft

2.4 FORCE TERMS

* Load Factor

A crash force can be expressed as a multiple of the weight of an

object being accelerated. A crash load factor, .ilen multiplied by a

weight, produces a force which can be used to establish ultimate

static strength (see StatiL Strength). Load factor is expressed in

units of G.

* Forward Load

Loading in a direction toward the nose of the aircraft, parallel to th6 aircr.t longitu•d"nal ro"A axs

a Aftward Load

Loading in a direction toward the tail of the aircraft, parallel to

the aircraft loý'oicudinal (roll) axis.

• Downward Load

loading in a downward direction parallel to the vertical (yaw) axis

of the aircraft.

0 Upward Load

Loading in an upward direction parallel to the vertical (yaw) axis

of the aircraft.

* Lateral Load

Loading in a direction parallel to the lateral (pitch) axis of the

airrraft.

W Combined Load

Loading consisting of components in more than one of the directions

described in Section 2.1.

13
0 Crash Force Resultant

The geometric sum of horizontal and vertical crash forces: horizontal and vertical velociLy components at impact, and horizontal and

vertical stoppinq distances. The crash force resultant is fully

defined by determination of both its magnitude and its direction.

The algebraic sign of the resultant crash force angle is positive

when the line of action of the resultant is Tbove the horizontal,

and negative if the line of action is below the horizontal.

* Crash Force Angle

The angle between the resultant crash force and the longitudinal

axis of the aircraft. For impacts with little lateral component of

force, the crash force angle is the algebraic sum of the crash force

resultant angle plus the aircraft pitch angle.

AIRCRAFT LONGITUDINAL AXIS 7

S. ... 'IHORIZONT AL •

AIRCRAFT

PITCH ANGLE - /

CRASH FORCr /

ANGLE CRASH FORCE Ai'I.E =

/" RESULTANT ANGLE +

/ PITCH ANGLE

RESUL.TANT GVCTICAL

CRASH FORCE

/--1 .... RESULTANT ANGLE.

GilumIZONTAtL "

2.5 DYNAMICS TERMS

* Rebound

Rapid return toward the original position upon release or rapid

reduction of the deforming load, usually associated with elastic

deformation.

* Dynamic Overshoot

The amplification of decelerative force on cargo or personnel above

the floor input decelerative force (ratio of output to input). This

amplification is a result of the dynamic response of the system.

14
a 'fi ran smi s_ si b i 1 i_ -

The amplification of a steady-state vibrational input amplitude

(ratio of output to input). Transmissibilities maximie at resonant

frequencies and may increase acceleration amplitude similar to

dynamic overshoot.

2.6 CRASH SURVIVABILITY TERMS

* Survivable Accident

An accident in which the forces transmitted to the occup-int through

the seat and restraint system do not exceed the limits of human

tolerance to abrupt accelerations and in which the structure in te•

occupant's immediate environment remains substantially intact tu the

extent that a livable volume is provi ced for the occupants throughout the crash sequence.

* Survival Jnvelyoe

The range of impact conditio.s, including magnitide and directior of

pulses and duration of forces occurrirg in an aircraft accident,

wherein the occupiable area of the aircraft )erains substantially

intc;t, bulh during and following the impact: and the I-orces transmitted to the occupants do not exceea the 1nits of human tolerance

when current state-of-t'-,e-art r2-traint systems are used.

It should be ;noted that, where 'he occupiable volume is altered appreciably through elastic defo-mation during the impact phase, survivable ronditioirs may not ,ave existed in an accident that, from

postcrash inspection, outwardly appeared to be survivable.

a Strike Envelope

The extent of space surrounding a restrained occupant defined by the

flailing of extended body parts during a crash impact of the aircraft. Parts of the body may strike objects located within this

envelope.

2.7 OCCUPANT-RELATED TERMS

* Human Body Coordinates

In order to minimize the confusion sometimes created by the terminology used to describe the directions of forces applied to the body, a

group of NATO scientists compiled the accelerative terminology table

of equivalents shown in Figure 5 (Reference 21). Terminology used

throughout this guide is compatible with the NATO terms as

illustrated.

15
Headward

(+G ) Direction of

2 accelerative force

Vertical

Back to chest Headward - Eyeballs-down act o Laterst tiTailward - Eyeballs-up (sternumward Lateral right

(+G ) Transverse

Lateral right - Eyeballsleft

Lateral left - Eyeballs- right

Back to chest - Eyeballsin

Lateral 

La ltef left er Chest to back - out Eyeballs-

(-Gy) to back

(spineward) Note:

Tailward (-G The accelerative force on

(-Gz) the body acts in the same direction as the arrows.

FIGURE 5. TERMINOLOGY FOR DIRECTIONS OF FORCES ON THE BODY.

6 Anthropomorphic Dummy

A device designed and fabricated to represent not only the appearance of humans but also the mass distribution, joint locations,

motions, geometrical similarities (such as flesh thickness and load/

deflection properties), and relevant skeletal configirations (such as

iliac crests, ischial tuberosities, rib cages, etc). Attempts are

also made to simulate human response of major structural assemblages

such as thorax, spinal column, neck, etc. The dummy is strapped

into seats or litters and used to simulate a human occupant in

dynamic tests.

0 Human Tolerance

For the purposes of this document, human tolerance is defined as a

selected array of parameters that describe a condition of decelerative loading for which it is believed there is a reasonable probability for survival without major injury. As used in this volume,

designing for the limits of human tolerance refers to providing

design features that will maintain these condiLiuns at or below

their tolerable levels to enable the occupant to survive the given

crash impact conditions.

16
Obviously, the tolerance of the human body to crash impact conditions is a function of many variables including the unique characteristics of the individual person as well as the loading variables.

The loads applied to the body include decelerative loads imposed by

seats and restraint systems as well as localized forces due to impact with surrounding structures. Tolerable magnitudes of the decelerative loads depend on the direction of the load. the orientation

of the body, and the means of applying the load. For example, thle

critical nature of loads parallel to the occupant's spine manifests

itself in any of a number of spinal fractures, but typically the

fracture is an anterior wedge, or compressive failure of the front

surface of a vertebra. Forces perpendicular to the occupant's spine

can produce spinal fracture through shear failures or from hyperflexion resulting, for example, from jackknife bending over a

lap-belt-only restraint. The lap belt might inflict injuries to the

internal organs if it. is not retained on the pelvic girdle but is

allowed to exert its force above the iliac crests in the soft

stomach region. Excessive rotational or linear acceleration of the

head can produce corcussion. Further, skull fracture can result

from localized impact with surrounding structure. Therefore, tolerarnce is a function of the method of occupant- restraint as well as

the characteristics of the specific occupant. Refer to Volume II

for a more detailed discussion of human tolerance.

Rotation of the hips under and about the lap belt as a result of a

forward inertial load exerted by deceleration of the thighs and

lower legs, accompanied by lap belt slippage up and over the iliac

crests. Lap belt slippage up and over the iliac crests can be a

direct result of the upward pull of the shoulder harness straps at

the middle of the lap belt.

* Effective Weight

The portion of occupant weight supported b~y the seat with the occupant seated in a normal flight position. Since the weight of the

feet, lower legs, and part of the thighs is carried directly by the

floor through the feet, this is considered to be 80 percent of the

occupant weight plus thp weight of the helmet and any equipment worn

on the torso. Clothing, except for boots, is included in the occupant weight.

* Ili-ac Crest Bone

The upper, anterior portion of the pelvic (hip) bone. ihese

"invertLed saddle" bones are spaceid lateralliy about 1 fi, apart; the

lower abdomen rests between these crest bones.

Lai Bcl TL*iedown Strap~a so Negative-C Str.ýj Crot~ch $tra~p)

Strap used to prevent the tenisile force in shoulder strapsý from

pull ing the lap belt up when the rcstrained subject is exposed to

_Gx (cyeballs-out) acceleration.

17
2.8 SEATING GEOMETRY (See Eiqure 61

V1I',A14't. FIRON4 rTr75(N yr.Y POSITION. TO VtRTI(AL P2.AN7. Of

.JrUThIAI S.EAT P PFI.RFHCI MINI FORi VARiu FA AT

flACK AN(11L7

SFAT FIACKt AN4(,1.L (I-Ci -w154 (D slJ1IS (I XI 

10 '7.7 -9 .

101 7.4

9.9 5 Kill

2. r, - NV

121 6.3

13 I.t

141 5.3 2-WAY STAl AI.JU52SIt 'll 4 -WAY 71 A7 AfrJIJS IMLNJ

___________________- IMiIlIZONtIAL / V,'1.51N LIM

F'037T10 % 

INI~

------ M 511f1 (, ~At, 7

v. ifl 3'

1 V 4 M -.- 1 AIli NI 

K III1C U1l3 11.2 -M HA-M

111.111.1 A141b M) tilRAI.l 

. IMUIUIWJ 7I.ILI' HE1

-, 1 Mill

lII. 111 I (A P. P1 1 3 1 II l 2 71 N1A I

I. P23775 TAHGFNT AIIOLE 3AlI.L B3E A 3iINMl.! UP 5w AN[D A MIA "1 .4171 (i3, 20, . v33 3157.tcor17.115 1321. Hm I1413 lqIPO )U" SHALTL Arpl~y.

2. TH IRENUML3105n SHOWN, APPLY To A rA3ITI L UTAII FL1IGHT STATION ANDT P03'ULATIutI 0q).IILII.

3. t2llu M 31.,"3 IA3ELJ ONl 13- SLAT VIIOLA AtCU2.r. PON7 FIXED. WTINO

AITtIGIIA?1, LESS THAI1' 73C5AT 13L11Cf APIULE 15 THL*T.31MADLE.

FIGURE 6. SEATING GLOMETRY. (REFERENCE 22)
0 Design Eye Position

A reference datum point based on the eye location that permits the

specified vision envelope required by M1I-STD-850 (Reference 23),

allows for slouch, and is the datum point from which the aircraft

station geomoetry is constructed. The design eye position is a fixed

point in the crew station, and remains constant for pilots of all

stature via appropriate seat adjustment.

* Horizontal Vision Line

A reference line passing through the desýign eye position parallel to

V the true horizontal in norma] cruise position.

* Back Tangent Iine

SP A straight 1line in the midpl aneý of the scat paSsing tangent to the

curvatures of' a seat occupant's back when leaning back and natural ly compressing the back cushion. The seat back tang(ent 1line is

positioned 33 in. behind the design eyc position nicasured along a

perpendicular to the scat back tangent 1line.

a buttock Reference Line

A line in the li dpl are of tF2 eatvr a a] ie to thu hoionzontal vi sion

1line and tangent to the lowermost natural protrusion (if a selected

Si ZC Of Occupant si ttiny on the sedt cushion.

0 NutrA,]Seat- ReferenceQ fuint- (.LLThR

1 he Intersection of' the back tangenit liine and the b)uttoc.k ref erence

l ine. 1 he seat geometry arid location aire based on the NS[U'. 1 he

NSRP i s set with the seat in the no i nal mid-po sit ion of 1.hce seat

adjustment ranye. lhis seat position will place theý 50th per-enitile,

(seated height) mar, with his cye in the desýign eye position.

* Buttocýk Refcr~i.ccQ Point

A point 5.7b in. forward of the seat referenuic point on the buttoc-k

reference line. This point defines the vertical and long itudinial

position of the approxi mate ho ttoms, of' the, i sch ial tui)Lros it ic u

thus representing the lowest poinits on the pel vic St)ructure and the(-

1)01nts that will support the must load dur rigy downiward vertic-al

loading.

a IeLL, LkLLUILIihie ref ererc- 1line parallel to the horiioiital vision line(- pa~ssitv'j

u ndcer the, t anyrig nt to the 1 owe st po ilit on tlie( hf eel i n the niorma l11

operat ional position, riot rIeC-sLQ.r ily co i n'(Jdtal i With the TI oUr

line.
2.9 STRUCTURAL TERMS

o Airframe Structural Crash Resistarce

The ability of an airframe structure to maintain a protective shell

around occupants during a crash and to minimize accelerations applied to the occupiable portion of the aircraft during crash impacts.

a Structural Integrity

"The ability of a structure to sustain crash loads without collapse,

failure, or deformation of sufficient magnitude to: (1) cause

injury to personnel or (2) prevent the structure from performinq as

intended.

* Jat th

The maximum static load which can be sustained hy a structure, often

expressed as a load factor in terms of G (see Load factor, Section 2.4). Also known as ultimate static load.

G Strain

The ratio of change in icngth to the original length of a loaded

component.

Deformation or fracture of structure to the point of loss of useful

load-carrying ability or useful volume.

Loss of Ioad-carrying capability, usually referringi to structural

linkage rupture or collapse.

L iit dLa

In a structure, l imit load refers tu the luad th'o struLture will

carry belore yielding. Similarly, in an eryergy-absurlirny duvi c:, it

represents th e load at which the devi cC defL) rm•. ir performi nrj it

function.

SL9Sad i LImII I .L L j_ ,itii 1)v iQe, PrL Luer_• A d_)_h' r

1 hse are interchanyeable names of devices u',cd t, 1 imit the load irl

a structure tu a prescl ectud vaILie. 1hisc devi -cs ab',orh enfrgy hy

providing a resistive inrce aipl iud over a deýformation distariLc

without siynif icant ela',tic rebound,

* fi , Lnerqy Absorbed (ýLA)

1hie en•ergy absorbed by a un,;rgy aburbirig dcvie uor structure

divided by its weight.

20
* Bottoming

The exhaustion of available stroking distance accompanied by an increase in force, e.g., a seat stroking in the vertical direction exhausts the avwilable distance and impacts the floor. With respect

to energy-absorbing structure, bottoming is a condition in which the

deforming structure or material becomes compacted and the load increases rapidly with very little increased deformation.

* Bulkhead

A structural partition extending upwards from the floor and dividing

the aircraft into separate compartments. Seats can be mounted to

bulkheads instead of the floor.

* Basic Structural Design Gross Weight (BSDGW4

The structural design gross weight is cited in the MIL-STD-1374 (Reference 24), Part I, "Group '-'eight Statement-Dime.tsional and Structural Data", and is further explained in the detail system

specification for the aircraft.

2.10 FIJEL, OIL, AND HYDRAULIC SYSTEiL TERMS

o ,Crash ,sistart IF uel Tank

A tank which conforms to MIL-T-27422 (Reference 25).

* Crash-Resistant Fuel Sst.m

A fuel system designed to conform to M!L-T-27422, MIL-STD-1290,

ADSiIB, arid other related specifications and standards.

* Frnugib Ie Attachment

An attachment uossessing a part that is designed to fail at a

predetermined location and/or load.

* U1adder Tank

A flexible fuel tank, usually contained or supported by other more

rigid structures.

* 1 ucl Pukmp

A pump installed in the fuIel syStem to move fuel . Usually located

at one or mnore of the following places: the tank, the engi ne, or

the interconnect iny plumbing.

* luel Valve

Any valve, other than a sC I f-seal iig breakaway valve, contained in

the fuel supply system, such as fuel shiutof f valves, check valves, etc.

21
* Self-Sealing Breakaway Valve

A valve, for installation in fluid-carrying lines or hoses, that

will separate at a predetermined load and seal at one or both halves

to prevent dangerous flarmmable fluid spillage.

2.11 IGNITION SOURCE CONTROL TERMS

a Fire Curtain

A baffle made of fire-resistant material that is used to prevent

spilled flammable fluids and/or flames from reaching ignition

sources or occupiable areas.

* Fire-Resistant Material

Material able to resist flame penetration for 5 min when subjected

to a 2000 OF flame and still be able to meet its intended function.

* Firewall

A partition capable of withstanding a 2000 OF flame over an area

of 5 sq in. for a period of 15 min without flame penetration.

S ,lamale rluid

Any fluid that ignites readily in air, such as hydrocarbon fuels and

lubricants.

* Flow Diverter

A physical barrier that interrupts or diverts the flow of a liquid.

a Igjnition Temperature

The lowest temperature at whirh a flammable mixture will ignite when

introduced into a specific set of circumstances.

*0 1ling

lhe rendering of an aircraft system or the atmosphere surrounding

the sysLem incapable of supporting combustion.

2.12 INTERIOR MATERIALS SELECTION TERMS

* Aiuoignition Temperature

The lowest temperature at which a flammable substance will ignite

without the application of an outside ignition source, such as

flames or sparks.

* Flame-Resiitant Material

Material that is self-extiniuishing after removal of a flame.

22
* Flashover

The sudden spread of flame throughout an area due to ignition of

combustible vapors that are heated to their flash pGint.

* Flash Point

The lowest temperature at which vapors above a combustible substance

ignite in air when exposed to flame.

* Intumescent Paint

A paint that swells and chars when exposed to flames.

* Optical Density (Ds)

The optical density is defined by the relationship

Ds = log -00

where T is the percent of light transmission through a medium (e.g.,

air, smoke, etc.).

2.13 DITCHING AND EMERGENCY ESCAPE TERMS

* Brightness

The luminous flux emitted per unit of emissive area as projected on

a plane normal to the line of sight. Measured in foot-lamberts.

* Candela (cd_

A unit of luminous intensity equal to 1/6n of the luminous intensity

of one square centimeter of a blackbody surface at the solidification temperature of platinum. Also called candle or new candle.

* Class A Exit

A door, hatch, canopy, or other exit closure intended primarily for

normal entry and exit.

a Class B Exit

A door, hatch, or other exit closure intended primarily for service

or logistic purposes (e.g., cargo hotches and rear loading ramps or

clamshell doors).

* Class C Exit

A window, door, hatch, or other exit closure intended primarily for

emergency evacuation.

23
9 Cockpit Enclosure

That portion of the airframe that encloses the pilot, copilot, or

other flight crew members. An aircraft may have multiple cockpits,

or the cockpit may be physically integrated with the troop/passenger

section.

* Ditching

The landing of an aircraft on water with the intention of abandoning

it.

* Emergency Lighting

Illumination required for emergency evacuation and rescue when

normal illumination is not available.

* Exit Closure

A window, door, hatch, canopy, or other device used to close, fill,

or occupy an exit opening.

* Exit Opening

An opening provided in aircraft structure to facilitate either

normal or emergency exit and entry.

a Exit Release Handle

The primary handle, lever, or latch used to open or jettison the

exit closure from the fuselage to permit emergency evacuation.

* Foot-candle (fcJ

A unit of illuminance on a surface that is everywhere one foot from

a uniform point source of light of one candela.

e Foot-lambert (fL)

A unit of photometric brightness or luminous intensity per unit

emissive area of a surface in a given direction. One foot-lambert

is equal to 1/n candela per square foot.

* Illumination

The luminous flux per unit area on an intercepting surface at any

given point. Measured in foot-candles.

24
3. AIRCRAFT DESIGN CRASH IMPACT CONDITIONS AND HUMAN TOLERANCE

3.1 INTRODUCTION

Design criteria for aircraft crash impact conditions and the response of the

human body to those conditions are presented in this chapter. Principles,

data, and analysis methods that influence the survivability of aircraft occupants in crash impact conditions are summarized. The reader is referred to

Volume II for a more complete discussion.

3.2 DESIGN CONDITIONS FOR IMPACT

3.2.1 Application

Although improvements in crash resistance can be achieved in existing aircraft by retrofit systems, such as energy-absorbing seats or crash-resistant

fuel systems, the improvements are limited and may result in prohibitive

weight and cost penalties if requirements are too severe. Retrofit decisions

are made as the result of trade-offs between the benefits in survivability

and the penalties of cost and weight. An aircraft should be designed as a

system to provide the required occupant protection for the recommended velocity changes because deceleration is a design variable, a function of the

structural stiffness of the fuselage, Consideration of crash resistance in

design of the complete aircraft system eliminates many of the limitations

inherent in retrofit and makes possible the design for more severe crash

impact conditions without significant weight penalties.

3.2.2 Deceleration Pulse Shape

Experimental data obtained in full-scale crash tests of helicopters, light

fixed-wing aircraft, and fixed-wing transports indicate that the deceleration

pulse shape for major impact in accidents can t,[ represented to a satisfactory degree for most engineering purposes by a triangle, as shown in Figure 4. Energy-absorbing landing gear on new aircraft will produce a lowerlevel deceleration plateau preceding the fuselage contact, thereby reducing

the energy that must be absorbed by fuselage crushing. However, the shape of

the deceleration pulse during fuselage contact with the ground will still

approximate a triangle.

3.2.3 Impacted Surface

Statistically, the crash surface most frequently impacted is sod. It is

recommended that sod with a California Bearing Ratio (CBR) of 2.5 be accepted

as the standard for crash-resistance design. Trees are the second most

frequently impacted obstacle; however, the secondary (in this case, major)

impact would still be with sod.

3.2.4 Velocity Change

Velocity changes for crash survival design purposes are specified by MIL-STD1290. Table 2 gives these velocity changes in feet per second. In addition

to these conditions, the designer should consider longitudinal, lateral, and

vertical impacts where the aircraft has pitch, yaw, and/or roll relative to

the flight path. Attitude angles are presented in Volume 1!.

25
TABLE 2. CRASH IMPACT DESIuN CONDITIONS, WITH LANDING

GEAR EXTENDED, NIL-STD-1290

Condition Impact Direction Object Viccity Change

No. (Aircraft Axes) Impact _AV (ft/sec)

1 Longitudinal 20

(cockpit) Rigid

2 Longitudinal vertical 40

(cabin) barriers

SVertical* Rigid 42

4 Lateral, Type I** horizontal 25

5 Lateral, Type 1 1**a surface 30

6 Combined high

angle* Rigid

Vertical horizontal 42

Longitudinal surface 27

7 Combined low

angle Plowed

Vertical Soil 14

Longitudinal 100

*For the case of retracted landing gear the seat and airframe combination shall have a vertical crash impact design velocity change

capability of at least 26 ft/sec.

**Type I - Light fixed-wing aircraft.

***Type II - Rotary-wing, including tilt-prop/rotor aircraft.

Note: See Volume IT for vehicle attitude.

3.3 HUMAN TOLERANCE TO IMPACT 4

3.3.1 General

Results of research on tolerance of the human body to impact forces are presented in Volume IT, Chapter 5. Although numerous experiments have been conducted and a wealth of information has been collected, very few criteria that

may be useful in system design have been developed and validated. in this

chapter, those criteria that are generally accepted for practical application

26
in assessing the crash resistance of an aircraft system are presented. As

discussed here, these criteria may be used to determine the acceptability of

an aircraft or components, such as seats and restraint systems, based on the

results of dynamic testing with anthropomorphic dummies or computer simulations, as discussed in Volume IV. Criteria are presented here only if validated quantitative values have been determined. Injuries to various body

parts are discussed in Volume II.

3.3.2 Whole-Body Tolerance

Tolerance of the human body to abrupt acceleration has been shown to depend

on the magnitude and duration of the applied force, as well as the direction

and rate of onset. Data presented by Eiband (Reference 26) for occupants

having upper torso restraint are summarized in Figures 7 and 8 for spineward

(-G.) acceleration and in Fiqures 9 and 10 for headward (+Gz) acceleration. Human tolerance to lateral (Gy) acceleration has not been extensively studied. However, based on the testing that has been conducted, a

maximum latere.1 acceleration of 20 G at a duration of 0.1 sec is suggested

for design.

An acceptable personnel restraint system for Army aircraft should include

upper torso restraint, regardless of seat orientation. However, for reference and for comparison with the above values, a spineward (-.G,) human toleranre level of 20 G and a lateral (G') level of 10 G are recommended for

lap-belt-only restraint. These levelý are based on experiments with human

subjects in which minor trauma were experienced.

Although Figures 7 through 10 indicate the regions of acceleration and rate

of onset that may be considered acceptable for the aircraft interior, they do

not permit complete evaluation of such protective systems as restraint systems, c-eergy-absorbing seats, or protective padding. Injury criteria for

critical body parts, such as the head and spinal column, must be employed in

order to answer such questions as whether a seat has sufficient stroking distance, or whether a given shoulder belt webbing has acceptable stiffness.

3.3.3 h'eaU Tlljury LIF I LtL

Various criteria have been used as predictors of head injury. Cuncussive

threshold values have been identified for four such criteria: peak G, peak

transmitted force, Severity Index, and Head Injury Criterion. The Severity

Index is oefined as

n S= J andt (1)

t

0

where SI = Severity Index

a = acceleration as function of time

n = weighting factor greater than 1

t = time

27
00

F-- La

ii Lai LAJ

'-CJ LLi-

'A LU

LLOLLI

I CIO

____n - -

II r23

1it, '-. C

h '02 T E.n - -

-4 0 04F 

> 040

0 r0 d)>D

'Io o

E9 631311-fA AO NOIiVUI31103V wAdojimnf

28
SUBJECT SUPPORT

o3 Human Lap, shoulder, thigh,

an d chest straps

0 Hiuman Lap, shoulder, thigh,

and chest straps

OHuman Ldp, shoulder, thigh,-

and chest straps ' Vchimpanzee Mlilitary lap and shoulder

straps

AChimpanzee 3-in, cotton webbing, 5 Acceleration

horizontal, 2 vertical

straps

60- - -

No sok N hc

7- 40 --- Cardiovascular shoc -shock .~ sin - -(conjun~ctival I llSigns and retinal -h - hck hemorrhage)

M:20---- ~

4 0f _ 1\ 0' - 44 4

::- 6 Onset rate max

FSTime

1-..4 0L1 J.L J L

Um

.001 .002 .04.006 .01. .02 .0 0 1 .2 .4 .6 1

TIME, SEC

FIGURE 8. INITIAL RATE OF CHANGE OF SPINEWARD ACCELERATION

ENDURED BY VARIOUS SUBJECTS. (FROM REFERENCE 26)

and the Head Injury Criterion (HEC) of Federal Motor Vehicle Safety Standard 208 is calculated according to

adhet strap

HIC = max (sf a(t)dt ( - t 1 ) (2)

(f TýI~mn L loldr tp thi*h

where a is the resultant head acceleration, and t, and t2 are any two

paints in time during the crash event.

29
4J4

E-

'-4i

L.0i

CDC~

_ _ _ _: LA-~

LA- C-)

- L C3)>

U--

c~o

01

0 01n1 

-44

0 0-0IIJ ~DOIf LONL.J1~l 

300
Sob~eLt Support

lomar = Lap antd shcuLdei straps

an Face curtain, lap and shoulder straps,

soft leather seat cushion

0 ias A'-rest (elbow only), lap and

shouldeir straps, pa.zachute, 2-in.

s::q:-:ri:,:r cushion

.D \ vuman Armtrest (elbow only), lap and I shoulder straps, wood hluck and :h pong.-ruLbz~r air cushuon Ilr~~

Z, Ch-Pa- - SUjnn; a cLdte-typ , restraint, Acc a

LU I-lud ln' lap and Ch bt St 1apSI

• 80

I " I

M

40

D- 10 - "_ - - - - 9A.. _

CU± !c; ,IC 1 1I / it. 1/

2 e'_im ~

.001 .002 .00O4 .006 .01 .02 .04 .01 .1 .2

TIMr_, SEC

FIGURE 10. INITIAL RATE OF CHANGE OF HEADWARD ACCELERATION

ENDURED BY VARIOUS SUBJECTS. (FROM REFERENCE 2•6)

Aircrewmen have experienced concussive head injury from helmeted head imnpacts that exceeded the following values for the four criteria: peak head accelerations that exceeded 150 G, peak force levels transnuiLted t.o the head that

exceeded 1500 Ib, Severity index values that exceeded 600, and Head In iury

Criterion values that exceeded 500. These values should be taken as the

limits of human tolerance to corncussior when using these (riteria as pre- dictors of head injury.

3.3.4 _Spina_£ I__niury Criter~ia

Although the Dynamic Response Index (DR]), as illustrated in Section 5.9.2 of

Volume II, is the only model correlated} ext~ensively for ejection seat spinal

31

U0 IU .00 U0 l . 2 . 4 O 1.
injury prediction, it has serious shortcomings for use in accident analysis.

It assumes the occupant to be well restrained and erect, so that the loading

is primarily compressive, with insignificant bending. Although such conditions may be assumed for ejection seats, they are less probable for helicopter crashes in which an occupant may be leaning to either side for better

visibility at the time of impact. Further, the DRI was correlated for ejection pulses of much longer duration than typical crash pulses.

A more detailed model of the spinal column would yield more realistic results, but injury criteria for the more complex responses have yet to be

developed. Consequently, the DRI is not. recommended as the critcrion for u,.e

in designing crash-resistant seats. Rather, the data presented in Figure 8

are recommended for use until more comprehensive data and criteria are

developed.

3.3.5 Leg Injury Cr-ti.ria

Femoral fracture due to longitudinal impact on the knee has been studied

extensively, probably because of the frequency of this type of injury in automobile accidents. A criterion that assesses the dependence of the perrnissible human knee load on the duration of the primary force exposure has been

suggested in Reference 27. The permissible peak knee load suggested for

design is given by

r =S0 - 160 t, t < 20 mseCI

F = 2000, t > 20 msec (3)

where F is in pounds and t in msec.

3.3,6 Tolerance of Other Body Parts

Although some research has been conducted on the tolerance of other body

parts, such as the neck, thorax, and abdomen, well-defined, valid criteria

have not been established. The results of this research are discussed in

Volume 1I, Chapter 5.

3.4 HUMAN BODY DIMENSIONS AND MASS DISTRIBUTIONS

3.4,1 General

Anthropometric measurements are external dimensions of the human body that

can be used to define aircraft requirements such as seat height and width,

eye height, or cabin height. A specialized type of anthropometric measure

ment is the "link length," or distance between joint centers, which can be

used in locating control poitions and is essential fur the design of matliematical or physical simulators of the human body. Finally, the inertial

properties of the body and parts of the body also are required in the design

of human simulators.

32
3.4.2 AnthroDometry

Two types of anthropometric measurements have been recorded, and the use of

both types in vehicle design has been summarized in Reference 28.

In the first type, conventional dimensions of the body with subjects in

rigid, standardized positions are easily obtained. Extensive collections of

such data are used in clothing design and may determine certain vehicle

design parameters including seat height and eye height. lhe anthropometric

data of greatest potential usefulness, illustrated in Figure 11, for U.S.

Army male aviators and soldiers of the 5th, 50th, and 95th percentiles are

presented in Tables 3 and 4, respectively. Complete data can be found in

References 29 and 30.

SITTING HE:IGHT

CA LYE HELIGHT

FUNCG IONAL REACH L' SHOULDER E.A.TH

S1-tOUI. DLk- ICJC)W LLEN'I H "I

ELBOW -FINGERTIP L1ENG1 11'I

BUTTOCK-(N[EL LENGIH KL LG1

POPLIT LAL HLIGiHT I H-'-4-IP BREADTH

BUT TOCK -POI'LIl [AL LHNGl H

FIGURE 11. CONVENTIONAL SEATED ANTHROPOMETRIC DIMENSIONS.

lhc second type of anthropometric data, which may be referred to as workspace

dimensions, is more difficult to obtain and can be applied only to the specific workspace studied. However, thfese workspace dimensions are essential

in designing aircraft interiors For maximum occupant protection.

Workspace dimensions must involve a consideration of body joints, the distance between them, and their rangy' of motion. Dermp)ster reported on an

extensive study of workspace requirements for seated operators, in which lie

determined "iink lengths" between effective joint centers for major body

parts (I1ferences 31 arid 32). lhcse link 1 eigths have a number u; (rash

resistance-related applications: (1) in developing or expanding the strike

33
TABLE 3. SUMMARY Of ANTHROPOMETRIC DATA FOR

U.S. ARMY MALE AVIAIORS (REFERENCE 26)

-Percentile (in.)

ttliurfment 5tLh O2Lt_ 9 St.h

We ight (1b) 133.0 P1.0 212.0

Stature 64.6 68.7 72.8

Seated height 33.7 35.8 37.9

Shoulder breadth 17.0 18.7 20 3

Functlonal reach 28.8 31.1 34.2

Hip breadth, sItting 13.2 14.8 16.7

Eye height, sitting 29.0 31.0 33.1

Vnee height, sitting 19.3 20.8 27.6 &

Elbow rest height, sitting 7.4 9.1 10.8

Puplitual height 11Wi 16.6 18.3

Shnulder-elhow length 13.3 14 4 11.6

Llfhaw f ingertip length 17.6 19.0 20.3

ButtoC.k-poplit aai length 17.7 19.3 21.0

buuLtock-knee lngth 22.0 23.7 25.4

TAbLL 4. SUMMARY Of ANIHROPOMLIRIC DATA FOR

FOR MALL SOLUILRS (RNiLLNLL 21)

Forcerit ic (in,.I

SMeasuremunt - t95tt

Weight (fb) 125.0 15 .0 202-0

Stature b4.5 66.7 73.1

Seated height 33.3 35.7 3b.1

Shoulder breedth 1b.3 17.6 19.6

Hip breadth, sitting 11.9 13.0 14.5

Lye height, sittil.. 28.0 31.0 33.3

KrT(i: height, sittilg 9.6, 21.3 23.1

Popliteal height lb.U 11 5 19.2

Shoulder-elbew length 13.3 14 5 15.7

Ll bw f iry:rtii : l, u i iii 17.4 10 U 20.4

Buttoek-pojrllLeal t 2lengtlh 1.0 1.6 21.3

buttock-knee length 21.0 23.4 25.3

34
envelopes shown in Chapter 6 of Volume II, (2) in designing crash test dumrmies, and (3) in providing numbers for mathematical simulators. Skeletal

joint locations and ranges of motion are presented in Section 7.2.3 of

Volume 11.

3.4..k Inertial PropertieS

Anthropometric dummies and mathematical simulations require inertial proper- ties of body segments, specifically moments of inertia, mass, and center-ofmass locations. Several studies of these properties have been made using live humar. subjects and cadavers, and such data as have been obtained should

be integrated into the design of any anthropometric dummy or mathematical

simulation. Results of several of these studies are summarized in

Reference 33.

3.5 CRASH TEST DUMMIES

All of the recently developed dummies were designed for automotive testing and are based on the anthropometry of a 50th-percentile U.S. civilian male.

In dynamic testing of an energy-absorbing seat, design for aircraft occupant weight can play a critical role. it would be desirable to evaluate a seat

for a range of occupant sizes. A 95th percentile dummy would verify the

strength of the seat structure and restraint system as well as the adequacy

of the energy-absorbing stroke. lestinn with a 5OLh-percent. ile dummy would

demn~s trdt;2 the perforliance of the system for an occupant of average height

and weight. A 5th-percentile dummy would p)rohably experience accelerations

of higher magnitude and would establish the severity of a given set of impact conditions for the smaller occupant. However, both the expense of dummy pur- chase arid the Lost of conducting dynamic tests may make such a test program impractical . An alternative procedure might be to establish the occupant pro- tection capability of a seat design by analysis and to conduct a dynamic test

with a 95th percentile dummy to verify system strength.

lhe desirin of differeimt anthropomorphic dumimies for mil itary testing must be

based on the miliitary aviator population. Body dimnensions, joint locations

n,,, m ass ri-,ý d - :4k t'---1bi ... t£ ; l10 .. -l ............. 01---2UI~t L -1 S 'I S{~ l Il -- - , IIIIU A d-, ... I 1(1 , .. _ . y (J t_-s - Iz Z. - Id - t I. dal riad

tors has been gynerated as a tri -serv iec data base for three-dimensional,

mathematical models and test dummies (Reference 31) . lhese dimensions are

available to de igjners for quidance in the desigri of dummies.

Another factor that must be considered in dummy selection for aircraft seat

test i n1 is that none of the dumnies described for aul.umotive testing have

been designed for accurate response to vertical impact. The spinal column,

which is a critical region of human tolerance to aircraft crash loading, has

been designed to simulate response to -G. loading, rather than the more critical -G. loading. lhe articulated dummies developied for car crash testing

are not suitable for vert ical impact tests, because they do not usualiy r,,pr -. sent spinal compressive stirfn nes arid the large number of corwncted masses

prevent the. evaluation of seat forces from the deceleration measured in the

dummy. Sarrailhe (Reference 35) proposes that the base of the test dummy be

as narrow as the load bearing part of a sealed human to ensure that the seat

and seat cushion receive representative loading, lie states that the spinal stiffness of the dummy could affect the, behavior of the seat.

35
In 1986 General Motors Corporation's Hybrid III test dummy was incorporated

into the Department of Transportation's specifications for Part 572 test

dummies (Reference 36). At present, manufacturers have t~he option of using

either the original Part 572 dummy or the Hybrid III dummy for compl~ance

testing; beginning September 1, 1991, the Hybrid III dummy will be used as

the exclusive means of determining a vehicle's conformance with the performance requirements of Federal Motor Vehicle Safety Standard 208. Therefore,

it appears that use of the Hybrid III dummy, modified to improve its simulation accuracy to impact loading in the 4GZ direction and sized to 5th-,

50th-, and 95th-percentile versions of the U.S. Ar-my aviator, would provide

the best available simulation and is, thorefore, the reconmmended approach.

36
4. AIRFRAME STRUCTURAL CRASH RESISTANCE

4.1 INTRODUCTION

Salient features required in the definition of a crash-resistant structure

are summarized in this chapter. The user is referred to Volume III for additional information concerning the criteria or their sources.

In a crash situation, the basic requirements for occupant survival of impact

hazards are:

e The maintenance of a protective structural envelope.

* The attentuation of impact forces to maintain survivable

acceleration conditions.

To achieve the desirable occupant survivable conditions, the following basic

design requirements must be considered as an integrated problem, and a prac- tical solution must be obtained. Such design requirements should be included

in new aircraft, and existing designs should be improved by incorporating

these features where possible.

* The basic structural envelope surrounding occupied areas must be

designed to maximize its energy absorption capacity.

* The structure that makes initial contact with the ground must be

designed to minimize the probability of earth gouging and scooping

of soil. This will minimize the acceleration and force levels to

which the structure is subjected.

* All items attached to the structure must, where possible, be retained in survivable crash conditions. These items include large

masses, such as transmissions, engines, and rotor systems; internal

cargo and on-board equipment racks; externally mounted components,

such as fuel tanks, wings, and external stores; and the empennage

aId landuing gear. 1I LIle past, sheuuding of lar-ge-tass i tenUs ds

been considered advantageous under crash impact conditions. This is

true from the viewpoint of reducing the energy content of the aircraft and, hence, the loads acting on the structure in resisting

aircraft postimpact motions. However, it is possible that penetration of occupied areas could occur, and during postimpact motions,

the aircraft could traverse shed objects, causing high loading on the

structure. It is, therefore, better to maintain a known mass if an

optimum acceleration profile is desired for occupant survival. Thus

mass retention and landing gear integrity are required for optimum

crash resistance and occupant survivable conditions.

* In the case of helicopters, certain areas of the cockpit and cabin

structure must be reinforced to withstand loads induced by blade

strikes, impacts with external objects such as trees, and rollover.

In addition, if overhead-mounted crash-resistant seats are used, the

deflection of the overhead structure relative to the floor must be

minimized.

37
s Unoccupied areas of structure, such as the underfloor, nose, and

tail areas, must be designed to deform in a controlled manner to

absorb as much energy as possible. Such deformation must be consistent with the safety requirements of other installed systems,

such as fuel cells or seats, and should not intrude into adjacent

occupied areas.

A crash can involve a wide range of dynamic conditions, from a simple unidirectional impact to a complex combination of rotational and multidirectional

impact conditions. The current requirements for Army light fixed- and

rotary-wing aircraft are summarized in Tables 5 and 6. Any light aircraft

designed to similar criteria would exhibit improvements in crash resistance.

A summary of desirable features for overall crash resistance is shown in

Figure 12 for a single-rotor helicopter. Similar features should be implemented in all designs, whether fixed- or rotary-wing, to provide survivable

conditions for all occupants.

When a more severe crash does occur, Lhe service life of the aircraft is usually ended, and the only structural requirement is to provide occupant protection. In order to provide such protection, the design must permit large

deflections of structural members and joints as well as loading in the plastic range of stress. Excessively strong airframe structure is no more acceptable than understrength structure for adequate crash resistance. Not only

will unnecessary strength result in an unacceptable weight penalty but on

impact high G levels that compromise occupant survivability may be generated.

TABLE 5. PERFORMANCE REQUIREMENTS UNDER CRASH IMPACT CONDIIIONS PER TABLE 2

Condition Impact Percentage Volume

No. Direction Reduction Other Requirements

1 Longitudinal No serious hazard to Does not impede postcrash

(Cockpit) pilct/copiloL. egress. Engine transmission, rotor system

remain intact and in place.

2 Longitudinal 15 maximum length re- Inward buckling of side

(Cabin) duction for passenger/ walls should not pose

troop compartment, hazard to occupants or

restrict their evacuation,

3 Vertical 15 maximum height re- G loads not injurious

duction in cockpit and

passenger/troop

compartment.

4 & 5 Lateral 15 maximum width Lateral collapse of occureduction. pied areas not hazardous,

no entrapment of limbs.

6 Combined No serious hazard to

High Angle occupant due to cockpit/

cabin reduction.

7 Combined No serious hazard to

Low Angle occupant.

38
TABLE 6. OTHER STRUCTURAL PERFORMANCE REQUIREMENTS

Velocity Vehicle Percentage

Impact Impacted Differential Attitude Volume Other

Direction Surface (ft/sec) Limits Reduction Requirements

Rollover Earth 90 D sideward or Minimal (door Forward fuselage buried to

1800 inverted hatches etc. depth of 2 in. (inverted

or any inter- assumed to be or on side). Load unimediate angle non-lodd carrying) formly distributed over

forward 25% of occupied

fuselage length. Can sustain 4 G without injury

to seated and restrained

occupants. All loading

directions between normal

and parallel to skin to be

considered.

Rollover (Post- Rigid Two 360u 15 maximum volume reimpact) Rolls (maximum) duction (5 percent

desired)

Earth Plowing Earth Preclude plowing when for-

& Scooping ward 25% of fuselage has

uniformly applied vertical

load of 10 G and rearward

load of 4 G or the ditching lnads of MIL-A-008865,

whichever is the greatest.

Landing Gear Rigid 20 10 0 Roll None. Plastic Aircraft deceleration at

+150 to -50 deformation of normal G.W. for impact

Pitch gear and mounting with no fuselage to ground

system allowable contact. All other A/C

structdral parts, except

blades, should be flightworthy following crash.

Landing Gear Sod 100 long.C* -50 Pitch 15 maximum volume No rollover, or if roll14 vertical ±100 Roll reduction (5 percent over occurs, two 3600

±20 Yaw desired) rolls without fuselage

crushing.

* Velocity at impact, not differential.

39
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0 40
4.2 AIRFRAME CRASH RESISTANCE

The aircraft structure should provide a protective shell for vehicle occupants

in crashes; moreover, the structure should allow deformation in a controlled,

predictable manner so that forces imposed upon the occupant will be minimized

while still maintaining the protective shell. In structural areas where large

structural deformations are anticipated, joints and attachments should be

designed to withstand large angular deflections and/or large linear displacements without failure. All exterior surfaces and all structures which could

be exposed to contact with the impact surface should be constructed of materials which characteristically resist sparking as a consequence of abrasion.

Unless otherwvise stated herein, the aircraft design gross weight (0GW) should

be used for the vehicle weight in the analysis described below. Directions

are assumed with respect to the aircraft (Figure 3) unless otherwise stated.

4.2.1 Long.itudinal Impact

4.2.1.1 Impact Conditions. The basic airframe should be capable of impacting longitudinally into a rigid abutment or wall at a contact velocity of

20 ft/sec without crushing the pilot and copilot stations to an extent which

would either preclude pilot and copilot evacuation of the aircraft or otherwise be hazardous to the life of the aircraft occupants. For such an impactr

the engine(s), transmission, and rotor system for helicopters should remain

intact and in place in the aircraft excpnt for damage to the rotor blades.

The basic airfirame of passenger-carrying helicopters should be capable of impacting longitudinally into a rigid abutment or wall at a contact vcolocity of

40 ft/sec without reducing the length of the passehlger/troop compartment by

more than 15 percent. Any consequent inward buckling of walls, floor, and/or

roof should not be hazardous to the occupants and/or restrict their evacuation. The aircraft should also be designed to withstand impact as in a lowangle missed approach. This impact in plowed soil (Figure 13) can result in a

rollover and side impacts which may crush and/or separate the fuselage. The

volume of the cockpit or the occupied passenger/troop compartment should not

be reduced by more than 15 percent (5 percent desired).

cko - ~ -- ~ ul L d uo !arrf fver, the Fuselage container should maintain structural integrity for a minimum of two 360-degree rolls. The static loads to be

considered for rollover analysis are described in Section 4.2.4.

4.2.1.2 Earth Scoopg. Design features for reducing the earth scooping

effects encountered in longitudinal impacts should include the following:

* Provide a large, relatively flat surface in those areas which could

gouge or plow, thereby increasing the aircraft's tendency to slide

over the impact terrain.

0 Minimize inward buckling of the fuselage nose or engine nacelle to

maintain skid surface integrity.

0 Design the nose section to preclude ans earth plowing and scooping

tendency when the forward 25 percent of the fuselage has a uniformly

appliedilocal upward load of 10 G and an aft load of 4 G. as shown in

Figure 14.

41
IMPACT CONDITIONS

1. SOIL OF CALIFORNIA BEARING RATIO=2.5

2. AIRCRAFT PiTCH (0)= 5' NOSE DOWN

3. AIRCRAFT ROLL (M)=±10'

4. AIRCRAFT YAW (3')=±20°

5. FLIGHT PATH ANGLE (a)=-8° DOWN

6. GROUND IMPACT SPEED = 100 FT/SF.C

7. iMPACT SINK SPEED =14 Fl iSEC

x

X GROUND LEVEL

FIGURE 13. LOW-ANGLE IMPACT DESIGN CONDITIONS (SIMULATED APPROACH

WITH ANlITORQUE LOSS UNDER POOR VISIBILITY).

4 G

10 G

FIGURE 14. NOSE SECTION DESIGN CONDITIONS.

4 Z
4.2.1.3 Fuselage Deformation. To minimize hazards to personnel created

by buckling or other deformation of the structur3, the aircraft should be designed to:

a Provide sufficient structural strength in the protective shell around

the occupants to prevent bending or buckling failure of the fuselage,

in accord with Table 5.

* Buckle the fuselage outward, if at all possible, rather than inward

into living space, when its collapse strength has been exceeded.

* Have the cargo restraints be effective even when fuselage bending

failure occurs.

4.2.1.4 Floor and Bulkhead. The floor structure should possess sufficient

strength and ductility to carry, without failure, loads applied by the occupants and cargo restraint systems even when deformation and substructure crushing occur. Consideration should be made for the specific loads and moments

applied by these -items to the suppcrting structure in the warped conditions

described in Chapter 5.

4.2.2 Vertical Impact

4.2.2.1 Impact Conditions. With the landing gear extended and the rotor/

wing lift equal to DGW, the aircraft should withstand a 42-ft/sec vertical

impact without reducing the height of the cockpit or cabin by more than 15 percent and/or causing the occupants to experience injurious accelerative loading. For this analysis, the aircraft attitude should be within +15/-5 degrees

of pitch and ±10 degrees of roll in accordance with MIL-STD-1290.

4.2.2.2 Design Application. Design applications for accomplishing the

above goal should include the following:

a Locate high-mass items so that they will not intrude into occupied

areas during the crash.

rI I dC .uf .. t 1r . cul ..... rus nng s..trength to ..... mo than

15 percent crush.

* Provide load-limiting structure beneath the floor.

* Provide load-limiting landing gear.

* Provide load-limiting seating for all occupants.

4.2.3 Lateral ImDact

Light fixed- and rotary-wing aircraft should withstand a lateral impact of 25

and 30 ft/sec. respectively, into a rigid barrier without reducing the width

of occupied areas by more than 15 percent. The deiign of the vehicle should

minimize the chance of the occupant being trapped betwee, the structure and an

impacting surface following failure of doors, canopies, or hatches.

43
4.2.4 Rollover Impact

The aircraft should be designed to resist an earth impact loading as occurs

when the aircraft impacts the ground and rolls to a 90-degree (sideward) or

180-degree (inverted) attitude. A rollover should not cause structural failure

or major intrusion into occupied areas. It should be assumed that the forward

fuselage roof is buried in soil to a depth of 2.0 in. for the inverted attitude

and that the load is uniformly distributed over the forward 25 percent of the

occupied fuselage length. It should also be assumed that the forward fuselage

side is buried in soil to a depth of 2.0 in. for the sideward attitude and that

the load is uniformly distributed over the forward 25 percent of the occupied

fuselagc length. The fuselage should be capable of sustaining a 4-G (i.e., 4.0

x aircraft DGW) load applied over the area(s) described for either the inverted

or sideward attitudes shown in Figures 15 and 16, respectively, without structural failure or more than 15 percent loss of living space. For both cases in

Figures 15 and 16, the 4-G distributed load should be analyzed for any angle of

load application ranging from perpendicular to the fuselage skin (i.e., compressive loading) to parallel to the fuselage skin (i.e., shear loading). When

designing for this condition, it should be assumed that all doors, hatches,

transparencies, and similar openings cannot carry any loading and that rotor

masts, wings, and tail boom are intact. These design conditions assume that

the aircraft becomes inverted after impact. They are not intended to provide

protection in an inverted impact.

4

f • •2-in. depth

of terrain (i.e. either

earth or water)

(parallel xy4 to x

x-y i'lane)•

1. 1vessure ont zoof to bh aptI1 i'd 1

uniforinly over -2eua rouultini(j 4 G_

from imlmersion int soil or water (telr"t-nd ic)lr to x-y i1aite toa -in. depth. • clrt - ln

2. 1. is thLe total length of fusela leo__

without either main or tail rotor t

bl.Adus. Shioded area is structural area

over which 4-(- load is applied

FIGURE 15. ROLLOVER, ROOF IMPACT DESIGN CONDITION.

44
141

4 G_ (erpendicu]ar

to x-z plane)

4 G,<2 (parallel to x-z plane)

Shaded area is

structural area

over which 4-,

load is applied

Tranor water

NI)TE::-•--•--•--•- • - 2-in. depth

]. Pressure on sides to be applied uniformly over area

ivýulLiIvj- fiu, iu.w ,L • ji J0 ! oil ] or witer to a 2 -in. r'Iptn.

2. 1, is the total length ot fuselaye without either main

or tail rotor blades.

FIGURE 16. ROLLOVER, SIDE IMPACT DESIGN CONDITION.

4.2.5 Wings and Empennage

As discussed in Section 4.1, the wings and empennage structure should remain

attached during a moderate crash. However, for fixed-wing aircraft, wing design should possess frangible characteristics to allow wings to break free

from the fuselage under high longitudinal inertia loads. This would limit

distributed impact loads caused by striking a barrier such as an earth mound.

Empennate structure should also be designed to collapse or break away during a

severe longitudinal crash impact. The structures should be designed to ensure

that failure occurs outside the occupant-protecting section of the fuselage.

For rotary-wing aircraft, wings used to support external stores prevent rollover in many accidents and should not be frangible, but should allow the

stores to separate under high-G loads while maintaining the structural integrity of the wing. However, the wing should break off before the fuselage

structure itself collapses.

The adjusted position of control surfaces such as flaps should not block doors

or other escape routes from the aircraft.

4.2.6 Engine/Transmission Mounts

For light fixed-wing aircraft, engine mounts should be designed to keep the

engine attached to the basic supporting structure under the crash conditions

45
cited in Table 2, even though considerable distortion cf the mounts and

support structure occurs. The basic structure supporting the engine should

fail or separate before engine mount failure occurs.

On helicopters, the transmission, rotor mast, rotor hub, and rotor blades

should not displace in a manner hazardous to the occupants during the following impact conditions:

* Rollover about the vehicle's roll or pitch axis on sod.

a Advancing and retreating blade obstacle strikes that occur within the

outer 10 percent of blade span, assuming the obstacle to be an

8-in.-diameter rigid cylinder.

Unless otherwise specified, all engines, transmissions, rotor masts, armament

systems, external stores, and rotor hubs that could be hazardous to the occupants should be designed to withstand the following ultimate load factors (G)

in the directions that cause those hazards and remain restrained:

* Applied Separately

Longitudinal ±20

Vertical +20/-10

Lateral ±18

* Applied Simultaneously

Design Conditions

1 2 _ 3

Longitudinal +20 +10 +10

Vertical +10/-5 +201-10 +10/-5

Lateral +9 +9 +18

4.2.7 Landing Gear

The landing gear should be capable of ground taxi, towing, grvond handling,

takeoff and landing roll, and landings including autorotative landings at

design sink speeds in accordance with AMCP 706-201 (Reference 37). Unless

otherwise specified, strength and rigidity requirements should be provided in

accordance with MIL-S-8698 (Reference 38). An analytical casting factor of A

1.25 should be applied for the design of all castings which will not be static

tested to failure, or which are not procured to MIL-A-21180 (Reference 39).

The yield factor of safety should be 1.0.

4.2.7.1 Landing Gear Location. The landing gear subsystem location should

minimize the possibility that a part of the gear or support structure will be

driven into an occupiable section of the aircraft, or into a region containing

a flammable fluid tank or line, in any accident falling within the crash conditions of Table 6. If this cannot be accomplished by location, the gear

should be designed to break away under longitudinal impact conditions, with

points of failure located so that damage to critical areas is minimized.

46
Failure of the landing gear should not result in a failure of any personnel

seat/restraint system or seat/restraint system tiedown. Also, failure of the

landing gear should not result in blockage of a door or other escape route or

prevent the opening of any door or other escape route.

4.2.7.2 Vertical Crash Force Attenuation. The landing gear should be of

the load-limiting type and should be capable of decelerating the aircraft at

DGW from a vertical impact velocity of 20 ft/sec onto a level, rigid surface

without allowing contact of the fuselage proper with the ground. Plastic deformation and damage of the gear and mounting system are acceptable in meeting

this requirement. The aircraft should be capable of meeting this requirement

in accidents with simultaneous fuselage angular alignment of ±10 degrees of

roll and +15 to -5 degrees of pitch. For the case of retracted landing gear,

the seat and airframe combination should have a vertical crash impact design

velocity change capability of at least 26 ft/sec. The landing gear should

provide energy absorption at sink rates up to 42 ft/sec onto an impact surface

within +10 degrees roll and +15 to -5 degrees pitch.

4.3 ANCILLARY EQUIPMENT RETENTION

Ancillary equipment is all removable equipment carried inside the aircraft

that could constitute a hazard if unrestrained during a crash. Typical items

are:

* Emergency Equipment 0 Aircraft Subcomponents

Oxygen bottles Panel-type consoltes containining

Fire extinguishers control circuitry

First-aid kits Radio and electrGnic equipment

Portable searchlights Auxiliary power units

Crash axes Batteries

Special Equipment

* Survival Equipment * Miscellaneous EQuipment

Survival kits Navigation kits

Life rafts Briefcases

Life jackets Log books

Locator beacons Flashlights

Special clothing Luggage

Food and water Toolboxes.

All ancillary equipment frequently carried aboard an aircraft should be provided with integrated restraint devices to ensure retention of the equipment

during any survivable crash of the severity cited in Table 2. Stowage space

should be provided for nonrestrained items that are not regularly carried.

This space should be located so that the items stored in it cannot become

hazards in a survivable crash. Stowage under energy-absorbing seats is not

acceptable.

4.3.1 Strength

Restraint devices and supporting structure for ancillary equipment should be

designed for static loads of 50 G downward, 10 G upward, 35 G forward, 15 G

aftward, and 25 G sideward. Load-limiting devices are recommended for

47
restraint of heavier equipment. However, load-limiter stroking should not

allow equipment to enter an occupant strike envelope.

4.3.2 Emergency and Survival Equipment Stowage Location

Equipment should be: (3) located close to the primary crew chief station, if

applicable; (2) stowed in easy view of crew and passengers; and (3) easily and

reliably accessible in an emergency. Equipment should not be placed in areas

where cargo shifting or fuselage distortion will prevent or impair access to

it. Extreme temperatures, abrasion, and uncleanliness should be minimized.

4.3.3 Release of Emerqency and Survival Equipment

Retenition devices used to restrain emergency and survival equipment should be

capable of quick release without the use of tools by one person using one

V:nd. Release should be effected by a single motion actuating one device and

:hould i•ot require more than 5 sec from time of contact with the actuating device to the time when the equipment either falls free or is lifted free. If

equipinent is stowed in an enclosure, no more than 5 sec should be required for

orening the enclosure and removing the equipment. Aircraft attitude should not

a',versely affect release device operation. It should be possible to see the

latch position (open or closed) of the release device. The release device

actualini' handle should be of a color that contrasts with the surrounding area

and be easily discernible in poor light or smoky conditions.

4.1 INTERFACE OF OCCUPANT AND CARGO RETENTION SYSTEMS WITH AIRFRAME

Both seats and cargo tiedowns require structural attachments capable of withstar.ding the applied loads without failure or excessive deformation. Although

aJditional seat design and installation r-quirements are discussed in Chapter 5

uf tnis volume, there are several important points to be considered where structural interface occurs. For example, the basic floor structure should evenly

distribute loading to the underfloor frames and longitudinal members. All seat

and cargo attachment fittings should be attached through the floor to primary

underfloor structure, i.e., either the heavy, full-depth longitudinal beams, or

suhstantial underfloor frame elements. The elements should be compatible with

the types and magnitudes of crash loading applied by the seat or cargo attachments. This includes reaching the loads and moments applied by the seats or

cargo with deformed floor and bulkhead structure.

The tiodown points must be designed for the worst case combination of cargo

weight, center-of-gravity height above the floor, and G conditions during the

crash..

If energy absorbers are used for the seat or cargo attachments, the attachments and their fasteners should be designed to the limiting load condition,

considering the effects of angular displacement relative to the floor. To ensure structural integrity, all seat attachments must be designed to withstand

or attenuate computed maximum load, with consideration for bottoming, or exhausting of available stroke. In the case of tiedown rings, which usually are

rated to a certain load capability such as 5000 lb, the attachments and structures must he capable of withstanding the worst case, angled load without yielding. Although cargo tiedown energy absorbers may be used, if a choice exists

between energy-absorbing and nonenergy-absorbing tiedowns, the design criteria

must be for the worst case, which will likely be the nonabsorbing equipment.

48
Structure surrounding an energy-absorbing seat must be designed to allow clearance for seat operation. Elastic deformation should be added to the envelope

of seat stroke in determining the required clearance. If a well is provided

in the aircraft floor to allow additional stroking distance, at least a 2-in.

clearance should be maincained between the outer edges of the bucket and the

innermost hardware extension on the sides or front of the well, including the

tracks.

4.5 OCCUPANT RETENTION

Seating and litter systems should ensure that occupants are retained in their

precrash positions within the aircraft. Seating and litter systems design

should be coordinated and interfaced with the design of the other surrounding

aircraft areas to achieve a completely integrated and efficient crashresistant aircraft system design. Seat and litter design should provide the

greatest practical amount of support and contact area for the occupants in the

directions of the most severe and likely impacts. Seats should provide an

integral means of crash force attenuation. Occupant comfort should not be

compromised to the extent that flight safety and/or crew efficiency is

adversely affected. Volume IV contains a detailed discussion of occupant

retention.

4.6 CARGO RETENTION

Cargo restraint should:

* Be as light as possible.

a Require minimum storage space.

* Be easy to install and remove.

* Be easily and reliably adjustable for different sizes and shapes of

cargo.

9 Provide sufficient restraint of cargo in all directions to prevent

injury to personnel in a survivable crash.

If the structure of the fuselage and floor is not strong enough to withstand

the longitudinal loads, load limiters should be used. Cargo restraints should

be capable of maintaining their integrity under longitudinal loads of 16 G

peak with a longitudinal velocity change of 43 ft/sec. Complete load and displacement requirements are presented in Table 7, and the requirements for the

longitudinal and lateral directions are illustrated in Figures 17 and 18. If

load-limiters are used, low-elongation restraining lines should be used to ensure the most efficient energy absorption.

Nets used to restrain small bulk cargo should be constructed of material with

low-elongation characteristics in order to reduce dynamic overshoot to a minimum. Restraining lines without load limiters used for large cargo, as defined in Table 8, for longitudinal restraint should b(. so arranged that maximum load in all lines is reached simultaneously, Restraining lines having

different elongation characteristics should not be used on the same piece of

cargo.

49
TABLE 7. CARGO RESTRAINT LOADS AND DIPSLACEMENT REQUIREMENTS

Luad Direction

Item (With Respect Restraint Controlled

No. to Floor) Load Displacement

1 Forward See Figure 17 See Figure 17

2 Aftward 5 G No Requirement

3 Lateral See Figure 18 See Figure 18

4 Downward 16 G No Requirement

5 Upward 5 G No Requirement

6 Forward See Figure 17 See Figure 17

and Combined

Lateral 4 G No Requirement

4.7 SPECIAL TILT-ROTOR CONSIDERATIONS

Crash resistance design considerations for tilt-rotor aircraft share many common items with conventional helicopter and fixed-wing aircraft. These include

the items listed in Table 1.

On conventional helicopters, the main rotor pylons arid engines generally are

located near the occupied areas, and adequate tiedown strength is required to

prevent potentially hazardous displacement of the large mass items into the

occinipd areas. nn tilt/rotor aircraft the pylons are located at the wing

tips, well away from the occupied areas. By allowing the wings to fail in a

controlled manner the aircraft mass is greatly reduced, and less material is

required in the fuselage structure to absorb the reduced aircraft kinetic

energy.

The aircraft design approach to control the wing, pylon, and rotor failure

modes is illustrated in Figure 19. By proper choice of the prop-rotor direction of rotation, the pylon and rotor are directed away from the occupied

areas in the event of a rotor ground strike. If a rotor ground strike occurs,

the composite blades typically exhibit benign failures and stay together because of the tensile strength of the fibers, even though severely damaged.

Although large blade sections are not expected to separate, debris from the

ground and rotor blade fragments may impact the fuselage sidewalls in the

vicinity of the tip path plane. The seating arrangements and the design of

the structure in those areas adjacent to the tip path plane should consider

this possibility.

50
Permissible

controlled

x. displacement

-td

floor:

Controlled

Aircraft •Ydisplacement flordevices

Net restraint Line restraint

20 - - __ _ T .. ....7I

Practical cargo displacement limit

18 (depending on aircraft) F

16

15

14 Acceptable failure area

(C 12

c• Mnimum acceptable IoAd curve

C) ample EFA

LL. 8 Restraint curves

.. ... : : . ...... ............ .... .... l . . .. .. ...... .. .. .......... :: .ii .!i i .i -;•. i !

S ailure load cur

2 ::---Unacceptable.performa. .. .. .. pne -................ ... ... *~~~. ccpab. perormnc ..- * ... ... ...... ... ..

00 1 5 10 15 20 25 30

CONTROLLED FORWARD CARGO DISPLACEMENT (x), IN.

FIGURE 17. LOAD-DISPLACEMENT REQUIREMENTS FOR ENERGY-ABSORBING

CARGO RESTRAINT SYSTEMS (FORWARD LOADING OF ROTARYAND FIXED-WING AIRCRAFT).

To maintain the fuselage occupied compartment area, the fuselage is designed

stronger than the win~gs. Similarly, the wings are designed stronger than the

pylon. Consequently, the pylon and wing will fail prior to the fuselage, pre- venting collapse of the occupant compartment (Reference 40).

51
O Failure load

----. Unacceptable performance

- Acceptable performance

M0inimum Practical cargo

S/-acceptable displacement

Co load curve limit (depending

c cceptable failure area

. 2

(A 0

0 5 10 15 20

CONTROLLED LATERAL CARGO DISPLACEMENT, IN.

FIGURE 18. CARGO LATERAL LOAD-DISPLACEMENT REQUIREMENTS.

TABLE 8. AIRCRAFT CARGO CATEGORIES

Small Bulk Cargo Large Rigid Cargo

(Net Restraint) (Line Restraint)

This class includes all boxes or This class inclues all rigid

unpacked cargo of approximately cargo of 3 ft3 or more in size.

3 ft 3 or less in size.

Examples: Examples: _

1. Ammunition boxes 1. Wheeled or tracked vehic.les

2. Foodstuffs 2. Aircraft engines

3. Medical supplies 3. Fuel barrels

4. Clerical supplies 4. Artillery pieces

5. Vehicle maintenance 5. Special weapons

components (priority cargo)

52
CONTROLLED WING FAILURE ALLOWS MASS RELIEF.

PREVENTS FUSELAGE COLLAPSE.

COMPOSITE BLADES HAVE

BENIGN FAILURE MODES

DIRECTION OF ROTATION CAUSES PYLONS/ROTORS

TO BE DRIVEN AWAY FROM FUSELAGE SHOULD

BLADE STRIKE OCCUR.

F!GIJRE 19. WING/ROTOR/PYLON CRASH FAILURE MODES.

4.8 TESTING

4.8.1 Aircraft System Testing

Instrumented, full-scale crash test(s) should be conducted: (1) to verify analyses performed, (2) to substantiate the capability of the aircraft system to

meet crash-resistance specifications, and (3) to gather further engineering

data on the impact response of aircraft structures. A more detailed discussion of full-scale aircraft system testing is presented in Volume ITI.

4.8.2 Landing Gear Crash ',iesting

Instrumented drop tests should be conducted: (1) to verify landing gear crash

force attenuation and crash loading strength characteristics analytically predicted and (2) to substantiate the capability of the aircraft landing gear to

meet the criteria of Section 4.2.7. Drop testing of wheel and skid landing

53
gear should be conducted in accordance with paragraph 9-2.3 of AMCP 706-203

(Reference 41). The 20-ft,/sec sink speed drop test should be conducted with

the landing gear oriented in a 5-degree nose down and 10-degree roll attitude

and drop tested onto a level, rigid surface with a sink speed of 20 ft/sec at

ground contact. Landing gear should also be drop tested in a O-degree roll,

pitch, and yaw attitude onto a level, rigid surface with a sink speed of

42 ft/sec at ground contact to demonstrate crash impact energy-absorption

capability. Rotor/wing lift for all drop tests should be equal to DGW.

Tests with a pair of gear mounted on an "iron bird" fixture simulate the

aircraft crash conditions more accurately than do tests on a single gear.

4.8.3 Cargo Restraint

Design loads are specified in Section 4.6. Static tests to these loads are

recommended. All deformation measurements are to be made at the floor level.

Sufficient dynamic tests should be made to assure that design predictions can

be accurately based on static test results.

4.8.4 Seat and Restraint System

Since proper performance of these items is critical for occupant survival, extensive qualification testing is required by MIL-S-58095. Testing requirements for seats and occupant retention systems are also described in

Volume IV.

4.8.5 Fuel System

Testing requirements for fuel systems are described in Volume V.

4,8.6 Ancillary Equipment Retention

Design loads are specified in Section 4.3, Static tests to these loads are

recommended. If applicable dynamic overshoot is likely, dynamic tests should

be conducted.

4.9gu:•u 1.... C o' EC KL ISTS

4.9.1 Landing Gear Design Checklist _

Yes No N/A

1. Will the gear withstand an impact velocity of up to

42 ft/sec without catastrophic failure?

2. Will the gear prevent the fuselage from contacting

the ground in a 20-ft/sec impact? - -

3. Will the gear survive a 10-ft/sec impact without

structural damage?

4. Will the gear remain attached to the fuselage after

impact?

54
Yes No N/A

5. Is the gear located to prevent penetration of occupied areas during the energy-absorbing stroke or in

the event of gear failure?

6. Has the gear been designed to absorb the maximum

energy consistent with available stroke?

7. Is the gear located to prevent rupture of fuel cells?

8. Is every blow-off valve located where fluid will be

confined or ejected outside the aircraft?

9. Has the gear been designed to avoid interference

with the stroke of energy-absorbing seats?

10. Has the gear been designed to continue to absorb

energy after initation of underfloor crushing?

4.9.2 Airframe Design Checklist

4.9.2.1 Fuselage

1. Are forward bulkheads canted aftwards below the

floor to prevent earth scooping?

2. Are the forward lower skin panels made of tough,

yet ductile, material to minimize tearing?

3. Are the forward lower skin panels shingled aftward

to prevent scooping?

4. Will the nose structure support an upward load of

10 G and an aftward load of 4 G applied over the

forward 25 percent of the fuselage without tailure

that would increase earth scooping tendencies?

5. Is the underfloor structure designed for energyabsorbing crush under upward loading while remaining

intact under longitudinal impact conditions?

6. Is structure designed to transfer loads due to

overhead masses to floor level without hazaraous

crushing of the occupied volume?

4.9.2.2 Wings and Empennaqe

1. Will the loss of wings occur in a manner that does

not endanger the occupants and that does not. destroy

the usable volume?

55
Yes No N/A

4.9.2.3 Rollover Structure

1. Will the forward fuselage roof support a 4-G load?

2. Are the side frame members designed for high load

capacity to prevent collapse during a rollover-type

impact?

4.9.2.4 Blade ImDact Protection

1. Are overhead longitudinal members extended continuously over cockpit areas?

2. Are upper surfaces smooth and is lateral structure

angled to deflect passing blades rather than allow

penetration?

4.9.2.5 Heavy Mass Support

1. Are the supports for massive overhead components

designed to withstand the following loads.

+18 G lateral?

+20 G longitudinal?

+201-10 G vertical?

2. ''ili the supports for massive overhead components

withstand the following combinations of loads:

+20 G long., +10/-5 G vert., j9 G lat.?

+10 G long., +20/-10 G vert., +9 G lat.?

.410 G long., +10/-5 G vert., +18 C lat.?

3. Do the engine mount. and fittin,]s, integral to the

engine as well as the aircraft structure, have

suFficient strength to remain intaci; until after

failure of major structural supporting members?

b6i
Yes No N__

4.9.2.6 Fuel Tank Installation

1. Are fuel tanks located above floor level and away

from possible impact surfaces?

2. Are fuel tanks located as far from occupiable

areas as reasonably possible?

3. Is fuel containment assured for all anticipated

survivable impacts?

4. Is the structure that supports fuel tanks smooth

and clean of projections to provide uniform support

and avoid puncture?

5. Are frangible and self-sealing couplings used in

fuel lines where relative displacements of structure may occur?

6. Are fuel tanks located outside the likely landing

gear motion envelope?

7. Have checklists of Chapter 6 been referred to for

fuel system design?

4.9.2.7 Seat and Cargo Installation

1. Is structure around seats designed to avoid interference with seat stroking and has sufficient clearance been allowed to enable efficient seat design

(see Volume III)?

2. Are seat and cargo attachment fittings secured

through the floor to primary structural members?

3. Are tiedown points designed for the worst case combination of cargo weight, center-of-gravity height

above the floor, and directions of loading aod

structural deflection?

4. Have checklists of Chapter 5 been referred to for

seat system design?

4.9.2.8 Emergenpc-ygess

1. Has the structure surrounding emergency exits been

designeo for minimum distortion?

2. Have the egress checklists of Chapter 6 been

referred to for emergency egress requirements?

57
5. AIRCRAFT SEATS, RESTRAINTS, LITTERS, AND COCKPIT/CABIN DELETHALIZATION

5.1 INTRODUCTION

This chapter summarizes the criteria for including crash resistance into the

design of aircraft subsystems that interface directly with the occupants.

These subsystems include restraint systems, seats, litters, cockpit controls,

and padding materials. The user is referred to Volume IV for additional

information concerning the criteria and their sources.

It is important to remember the basic operational difference between passenger seats and crewseats. The primary function of passenger seats and litters

is to provide a place for aircraft occupants to sit or lie during their transport, while the crewseats must provide the comfort, adjustments, and features

that aid crew members in accomplishing their operational responsibilities,

These functional requirements obviously are of highest priority; however,

crash resistance and the ability of the subsystems to help protect the occupant during crashes are also of extreme importance and can be accomplished

without significant degradation of comfort and operational aspects.

5.2 PRIMARY DESIGN CONSIDERATIONS

5.2.1 General

Occupant protection and survival in aircraft accidents shouid be a primary

consideratio in. III Udesign, LeveelUJmeIInL, and LtSeLin of aircr af seats and

litters. All operational requirements as specified in other design guides

should also be met. Adnquate occupant orotection requires that both seats

and litters be retained generally in their original positions within the

aircraft throughout any survivable accident. In addition, the seat should

provide an integral means of craoh load attenuation, and tha occupant's

strike envelope should be delethalized.

Several environmental and operational factors other than those associated

with crash resisttnce affect the design cf an adequate seating system. They

are very important in overall design and are discussed in Section 3.2 of

Volume IV.

5.2.L. Desin Conditions

The d'.sign impact con~litons for light fixed- and rotary-wing aircraft arc

presented in Voium, I and are repeated in Chapter 3, Table 2, of Volume i.

All seats, restraint systems, and litters should be designed to provide the

desired prfGrmarce in the design crash impact conditions. It must be remembered that, to produce a truly crash-resistant design, systems analyses must

consider likoly combinations of loadings, including potential losses of

energy-absorbing structure, such as lanaing gear.

5.2.3 Structurii Distortion

Structural distortion of the airframe and its resulting loading of the seat

must be considered in the design. A major consideration in providing crashresistant seating systems is the possibility of a local distortion in that

part of the aircraft to which the seat is attached.

58
In ceiling-mounted seats the efficiency of use of the available stroke distance must be considered. Energy-absorbing stroke should be provided to

maximize usage of the available space, but the effective stroke of a seat

considered to be rigidly attached (no energy absorbers between the seat and

ceiling) to the ceiling must be considered. The ceiling may deflect downward

at loads too low to make efficient use of the available stroke, a particular

concern for retrofit applications to older aircraft. A systems analysis

should be used to evaluate the advisability of using ceiling-mounted seats in

this situation and, if so, establish the correct combination of variables.

A considerable amount of the downward motion of an aircraft ceiling may be

elastic. It would be advantageous to eliminate from the occupant and

ceiling-suoported seat the rebound due to recovery of this elastic distortion. Consideration should be given to a device that allows vertical downward motion of the seat but restrains it from following the roof during its

elastic rebound.

Adequate support of the ceiling for the applied loads with low deflections

eliminates the problems mentioned above, and efficient use of ceiling-mouited

seats can be achieved in aircraft with such features.

Considerations for seats mounted on the floor, bulkhead or sidewall, including requirements necessary for the attachments to survive fuselage warpage, are presented in Section 5.4.5, Joint Deformation.

5.3 DESIGN PRINCIPrLES FuO StET ADU LITTERS

5.3.1 Seating System Orientation

There are several types of Army aircraft seating systems: pilot, copilot,

crew chief, gunner, observer, student, medical attendant, troop, and passenger. Cockpit seats are typically forward-facing; cabin seats may face in any

direction. Many are single-place seats, but in some aircraft two-, three-,

and four-occupant cabin seats are providpd. A single occupant se3t is the

preferred configuration in order to avoid situations in which the energyabsorbing systems of multi-unit seats are rendered ineffective due to partial

occupancy (insufficient weight to activate the energy-absorbing mechanisms at

loads within human tolerance limits). Seats should be interchangeable.

The rearward-facing seat is optimal for providing maximum support and contact

area in longitudinal impac's. The only critical impact sequence for the

rearward-facing seat is onE that involves a severe lateral component that

allows sideward movement of the occupant prior to application of the longitudinal or vertical Pulse. However, lateral torso movement can be minimized

by use of a torso restraint system of much lighter weight than that required

for other seat orientations. The rearward-facing cabin seat is preferred.

Those crew members required to face forward in the conduct of their duties

can be afforded adequate protection by the use of a restraint system consisting of shoulder straps, a lap belt, and a lap belt tiedown strap as discussed

in Section 5.7. Lap-belt-only restraint is undesirable, as noted in the

human tolerance section of Volume II. if all forward-facing passengers are

provided with adequate upper- and lower-torso restraint, forward-facing seats

dre acceptable as a second choice to -earward-facing seats. If a single,

59

Imm
diagonal upper-torso restraint is used, it should be placed over the outboard

shoulder of the occupant to provide restraint against lateral protrusion of

the occupant outside the aircraft or impact with the sidewall.

Previously, many side-facing seats were provided with lap belt restraint

only. Even with the addition of a shoulder harness or diagonal chest strap,

the tolerince to abrupt longitudinal acceleration is less than that for any

other orientation. The use of side-facing seats is least desirable for crash

safety; when no reasonable alternative exists, adequate torso restraint

should be provided. When a single, diagonal, upper-torso restraint is used,

it should be over the forward-facing shoulder (relative to the aircraft).

5.3.2 Litter Orientation

Litters should be installed laterally when practical, to provide more

positive restraint for expected combined crash forces. A lateral litter

orientation also will prevent detachment of the litter from its supports,

which may occur as explained in Reference 42. The litter should withstand

all of the conditions previously described for seats.

5.j.3 Materials

Designers should select materials that offer the best strength-to-weight

ratios while still maintaining sufficient ductility to prevent brittle

failures.

The degree of ductility needed in a seat's basic structural elements is

highly dependent upon whether the seat structure is designed to absorb energy

by the use of a separate load-limiting device or whether large plastic deflections of the basic structure are required. As a general rule, a value of

10-percent elongation is a rough dividing line between ductile and nonductile

materials. The 10-percent value is recommended as a minimum for use on all

critical structural members of nonload-limited seats because the exact peak

load is unpredictable due to pulse shape, dynamic response of the system, and

velocity change. A minimum elongation of 5 percent in the principal loading

direction is suggested for use on critical members of load-limited seats,

because the loads and strains are more predictable. Also, castings are not

recommended for use in primary structural load paths.

The effects of stress corrosion must be considered, as well as hydrogen

embrittlement, due to heat treating or various processing steps such as

pickling. In short, adherence to all the normal engineering design

principles is recommended.

Flammability and toxicity retardation requirements are discussed in Chapter 6. Upholstery padding and other materials used in seats should meet the

specified requirements.

5.4 STRUCTURAL CONNECTIONS

5.4.1 Bolted Connections

For the manufacture of basic aircraft structure, moss aircraft companies

recommend 15- and 25-percent margins of safety for shear and tensile bolts,

respectively. The margin of safety for shear and tensile bolts located in

60
load-limited portions of the seat, where loads can be predicted accurately,

can be reduced to 10 and 15 percent, respectively. Also, good aircraft engineering practice dictates that bolts less than 0.25 in. in diameter should

not be used in tensile applications because of the ease with whihl, these

smaller bolts can be overtorqued. Because of the obvious advantages of str"ucture being able to distort while maintaining load-carrying ability, fasteners

of maximum ductility for the application should always be selected. Where

possible, fasteners such as bolts and pins should have a minimum elongation

of 10 percent. A bolt loaded in shear should have a shank of sufficient

length to prevent application of the shear load on the threaded portion of

the bolt. For the best failure mode, bolts, pins, and joints shou'id be

designed to fail in bearing.

5.4.2 Riveted Connections

Guidelines for riveted joints are presented in MIL-HDBK-5, and it is recommended that these guidelines be followed (Reference 43).

5.4.3 Welded Connections

Welded joints can be completely acceptable and even superior to bolted or

riveted joints. However, strict inspection procedures should be used to

ensure that welded joints are of good quality. The cross-sectional area of

LMle basic niiaterial in the vitiituy of d welded joinL should be 10 percent.

greater than the area needed to sustain the design load. Welding processes

are discussed in Military Specifications MIL-W-8604, -6873, -45205, and

-8611; these specifications should be used as guides to ensure quality

welding.

5.4.4 Seat Attachment

Acceptable means of attaching seats to the cabin interior are listed below.

(Refer to Section 3.3.3 of Volume IV for a discussion of ceiling-mounted

seats.)

1, Suspended from the ceiling with energy absorbers and wall or

bulkhead stabilized.

2. Suspended from the ceiling with energy absorbers and floor

stabi:ized.

3. Wall or bulkhead mounted with energy absorbers.

4. Floor mounted with energy absorbers.

5. Ceiling and floor mounted (vertical energy absorbers above

and below seat).

Suspension or mounting provisions for all seats should not interfere with

rapid ingress or egress. Braces, legs, cables, straps, and other structures

should be designed to prevent snagging c- tripping. loops should not he

formed when the restraint system is in the unbuckled position. Cabin seats

must often be designed so that they may be quickly removed or folded and

secured. Tools should not be required for this operation. The time required

by one person to disconnect each single occupant seat should not exceed

61
20 sec. The time required by one person to disconnect multi-occupant seats

should not exceed 20 sec multiplied by the number of occupants. All foldable

seats should be capable of being folded, stowed, and secured or unstowed

quickly and easily by one person in a period not to exceed 2n sec multiplied

by the number of occupants.

5.4.5 Joint Deformation STo Drevent seat connection failures induced by fuselage distortion, structural joints should be capable of large angular displacements in all directions without failure. A floor-mounted seat designed properly for

structurally integral load limiting would also satisfactorily accommodate

floor buckling and warping under crash conditions. Figure 20 illustrates recommended limits of floor warping or buckling that should be withstood by

all floor-mounted seat designs. The mounts should be capable of withstanding a ±10-degree warp of the floor, as well as a ±10-degree rotation

about a roll axis of a single track. The angles are based on distortions

that have been noted in potentially survivable accidents.

The same general principles that apply for floor-mounted seats also apply for

bulkhead-mounted seaLs, except that the deflection and degree of warping of

the bulkhead appear to be less than those of the floor. A possible bulkhead

distortion configuration is shown in Figure 21. The recommended angular

deflection requirement for bulkhead-mounted seats is a 5-degree rotation in

the plane of the bulkhead. To accommodate local deformation, each attachment

of the seat to the bulkhead should be released to permit 010-degree rotations

in any direction.

Combined sidewall-mounted and floor-mounted seats require the same considerations as bulkhead-mounted seats. The sidewalls of aircraft tend to bow outboaid during impacts with high vertical loading. Therefore, it is advisable

that these seats be designed to accept relatively large distortions without

failure. Seats mounted to both the floor and the sidewall will require special design considerations. One way to provide the flexibility needed is to

include releases such as pin joints, oriented to allow rotation around an air- craft roll axis. An exapnle is shown in Figure 22. The attachments should

be designed to permit the angle 0 to reach 25 degrees at the maximum

dynamic deflection. Seats that are mounted totally on the sidewall should be less of a problem.

The underfloor, bulkhead, or sidewall structure must be designed to be compatible with the seat. For example, the design of structural releases between

the seat and the track may enable the seat to maintain its attachment during

large floor deformations but may add to the torsional loading on the underfloor beams. If a large downward load is applied to the floor structure

through a joint that does not carry moment (released), then the underfloor

beams must resist any moment that may be developed without assistance from

the seat structure.

5.4.6 Material

5.4.6.1 General. An elastic stress analysis, as ujsed in the design of

airframes and aircraft components subjected to normal flight loads, is inadequate for the study of all the structure in a crash situation. For normal

flight loads, keeping the stresses well below the material yield stress to

62
HYDRAULIC

CYLINDER

TEST FIXTURE STRUCTURE

TEST

STRUCTURE -' [ E C

DEFLECTED

POSITION

SEA - F 5- [0.087 RADIAN]

ORIGINAL

ATTACHMENTS--- • POSITION [0. S_ t • -/ - I 5o[0.087 A ,H

BULKHEAD

IX

Y

z

+100 [0.175 RADIAN]

POSITION • \

015R AD IAN -1O,[.o15to,7l os

I 1,0.17 RADIANj

FIGURE 20. FLOOR OR BULKHEAD WARPAGE REQUIREME4T FOR

STATIC LOADING OF SEAT. (REFERENCE 15)

63
-Bulkhead

,Seat attach

0 0 points (4 places)

Lower fuselage

(a) INITIAL CONFIGURATION

obutkcles

( ,---TCASCOFGRTN

FiGUE BKHEA 21 IN-LAN WARA5E

64c
Initial fuselage

configuration

Deflected

fuselage

configuration

4,4

2 5 max 0•.

FIGURE 22. PIN JOINT RELEASES ORIENTED TO ALLOW ROTATION

AROUND AN AIRCRAFT ROLL AXIS.

avoid permanent deformation is necessary because of fatigue problems and

other considerations. In a crash situation, however, where only one application of the maximum load is expected, fatigue is not a factor, and the

final appearance of a structural component or its subsequent operational use

need not be considered. Consequently, the load-carrying capacity of components deformed beycnd the elastic limit should be considered in determining

the ultimate seat strength. For certain items in the load path it is advisable to use the rupture strength as listed for many materials in MIL-HDBK-5

(Reference 43). The concepts of limit analysis or, in some circumstances,

large deformation analysis ma, be employed to make the best use of materials

in certain components.

5.4.6.2 Limit Analysis Concepts. Where ductile materials are used, strai,.

concentrations do not produce rupture prior to significant plastic deformation. If the geometric configuration of the structure permits only small

elastic deflections, a "rigid-plastic" mathematical model may be used. This

permits the use of a limit analysis, which assumes no deformation of structure until sufficient plastic hinges, plastic extensors, etc., exist to permit a geometrically admissible collapse mode.

65
Limit analysis is concerned with finding the critical load sufficient to

cause plastic collapse with the physical requirements of static equilibrium,

yield condition for the materials, and consistent geometry considerations.

Two useful principles are mentioned here: the upper and lower bound theorems. The upper bound theorem for the limit load (collapse load for a "rigid-plastic" structure) states that the load associated with the energy

dissipated in plastic deformation will form an upper bound for the limit

load. The lower bound theorem states that the load associated with a statically admissible stress distribution, which at no point exceeds the yield

conditions, forms a lower bound for the limit load. Use of the upper and

lower bound theorems to bracket the limit load for a given structure makes it

possible to obtain a realistic evaluation of the structure's load-carrying

capacity.

5.4.6.3 Large Deformation Analysis. It a structure contains elements that

will permit large, stable elastic deformations when under load, the equilibrium of the deformed state must be considered in evaluating ultimate

strength. For example, if a suitable attachment is made to a thin, flat

sheet rigidly fixed at the edges so as to load the sheet normal to the surface, a diaphragming action will occur. The equilibrium and stress-strain

(elastic-plastic) relations for the deformed state would determine the loadcarrying capacity. An example of this situation is a seat pan in which membrane rather than flexural stresses are important.

5.4.6.4 Strain Concentrations. Handbook stress concentration factors will

provide sufficiently accurate data to allow the designer to modify the structure in the vicinity of stress concentrations. When large deformations at

high load-carrying capacity are desired, as in energy-absorbing seats, these

areas frequently become strain concentration points, and rupture occurs due

to excessive strain in areas with little deformation and energy input. Large

amounts of energy can be absorbed in the structure only if large volumes of

material are strained uniformly.

5.4.7 Restraint System Anchorage

The seat designer should consider the effect of the anchoragp of the restraint system on the characteristics of the seat design. If possible, the

restraint system should be anchored to the seat rather than to basic aircraft

structure.

If the harness is anchored to basic aircraft structure, a desirable reduction

of loads on the seat frame results; however, the restraint system must be designed to permit the energy-absorbing deformation of the seat during an impact. For example, if a load-limited seat strokes vertically and the seat

belt is anchored to the floor, loosening of the belt permits the occupant to

either submarine or move laterally tinder the belt. When the harness is anchored to the seat structure, the problem of maintaining a tight harness is

reduced.

5.5 ENERGY-ABSORBING DEVICES

The seat structure, in order to perform its intended retention function,

should possess either (1) the capability of sustaining, without collapsing,

the maximum inertial forces imposed by the deceleration of the occupant and

66
the seat or (2) sufficient energy-absorption capacity to reduce the occupant's relative velocity to zero before structural failure occurs.* The

first alternative may result in an excessive strength requirement because

the input pulse shape and the restraint system and cushion elasticity can result in a large dynamic overshoot. Computer simulation and experimental investigation have shown that overshoot factors range from 1.2 to 2.0. This

would necessitate a seat design strength requirement of 24 G to 40 G to accommodate an input floor pulse of 20 G.

The second alternative of using collapse behavior (load limiting) appears to

offer the more practical approach to most seat design situations. With this

option, the seat structure would begin plastic deformation when the acceleration of the occupant and seat mass reaches a level corresponding to the critical structural load; the seat should absorb enough energy without failure to

stop the motion of the occupant relative to the aircraft. This energy should

be absorbed at force levels within human tolerance limits to provide the intended protective function. The energy can be absorbed either by plastic

deformation of basic aircraft structure or by the introduction of mechanical

load-limiting devices. Energy-absorbing motion of the seat can be provided

in all three directions as well as for all combinations of directions; however, it is absolutely necessary for the vertical direction. A properly

restrained occupant can withstand the loads associated with the design crash

impact conditions in the longitudinal (x) and lateral (y) directions but

cannot sustain the loads in the vertical (z) direction without injury. Therefore, the requirement for load reduction through use of energy-absorption

devices is mandatory for the vertical direction.

Energy-absorbing mechanisms in aircraft structures which transmit crash

forces to the occupant should stroke at loads tolerable to humans and should

provide stroke distances consistent with these loads and with the energy to

be absorbed. Desirable features of energy absorbers are as follows:

* The device should require a constant, predictable stroking force.

* The rapid loading rate expected in crashes should not cause unexpected changes in the force-versus-deformation characteristic

01 OleI uUVICC:.

0 The device should resist loads in the opposite direction to the

stroking (rebound) or be able to stroke in either direction.

* The assembly in which the device is used should have the ability to

sustain tension and compression. (This might be provided by one or

more energy absorbers, or by the basic structure itself.)

* The device should be as light and small as possible.

* The specific energy absorption (SEA) should be high.

* The device should be economical.

*The term "failure" implies a rupture of restraint linkage, while the term

"collapse" pertains to a state of active deformation with restraint integrity maintained.

67
a The device should be capable of being relied upon to perform satisfactorily throughout the life of the aircraft (a minimum of 10 years

or 8000 flight hours) without requiring maintenance.

* The device should be easily replaceable.

* The device should not be affected by vibration, dust, dirt, heat,

cold, or other environmental effects. It should be protected from

corrosion.

a The device(s) should decelerate the occupant in the most efficient

manner possible while maintaining the loading conditions within the

limits of human tolerance. A multiple-limit-load device, adjustable

for occupant weight, is desirable.

5.6 SEAT CUSH!ONS

5.6.1 General

The seat bottom and back cushions with which the occupant is in constant contact should be designed for comfort and durability. Sufficient cushion thickness of the appropriate material stiffness should be provided to preclude

body contact with the seat structure when subjected to either the specified

operational or crash loads. Seat bottoms made of fabric diaphragms should

have adequate clearance to prevent contact between the occupant and the seat

sVLucLUr'e ardi should be provided with means of Tightening to compensate for

sagging during use.

For seat cushions, the problem is one of developing a compromise design that

will provide both acceptable comfort and safety. The optimum aircraft seat

cushion should:

: Be lightweight.

0 Possess flotation capabilities.

0 Be nonflammable.

* Be nontoxic; not give off fumes when burned, charred,

or melted.

* Be tough and wear resistant.

a Be easily changeable.

I Provide comfort by distributing the load and reducing

or eliminating load concentrations.

6 Provide thermal comfort through ventilation.

a Provide little or no rebound under crash loading.

a Minimize motion during crash loading.

68
5.6.2 Requirements

For seats of light movable weight (less than 30 Ib), cushions should be comfortable with a maximum uncompressed thickness of 1-1/2 in., unless it can be

shown through analysis or through dynamic tests that the cushion design and

material properties produce a beneficial (reduced force transmissibility)

result.

For seats of greater movable weight, such as integrally armored seats, every

effort should be made to design a cushion that minimizes relative motion

between the occupant and the seat and that acts as a shock damper between the

occupant and the heavy seat mass. Again, dynamic analysis and/or testing

should be conducted to demonstrate that the cushion design produces a desirable system result over the operational and crash impact conditions of interest. In both lightweight and heavyweight seats the cushion should be as thin

as possible (without being uncomfortable) in order to minimize submarining.

5.6.3 Energy-Absorbing Cushions

The use of load-limiting cushions in lieu of load-limiting seats is undesirable. The only justifiable use of energy-absorbing cushions instead of loadlimited seats might be in retrofit circumstances where, because of limitations in existing aircraft, another alternative does not exist.

5.6.4 Net-Type Cushions

This type of cushion serves the same purpose as the filled cushion; however,

a net material is stretched over a contoured seat frame, and the body is

supported by diaphragm action in the net rather than by deformation of a compressible material. The net-type cushion might more properly be called a net

support. If a net support is used in the seat, its rebound characteristics

should be capable of limiting the return movement from the point of maximur,

deformation to 1-1/2 in. Net supports should not increase the probability of

occupant submarining or dynamic overshoot.

5.6.5 Seat Back Cushions

The back cushion should be of a lightweight foam material or net. The foam

can be a standard furniture type that meets the other requirements listed in

Section 5.6.2. Lumbar supports, particularly those that are adjustable by

the occupant, are desirable for comfort and because a firm lumbar support

that holds the lumbar spine forward increases the tolerance to iGZ loading.

5.6.6 Headrests

A headrest should be provided for occupant head/neck whiplash protection.

Headrest cushions are used only to cushion head impact and prevent whiplash

injury due to backward flexure of the neck. The cushioning effect can be pro

vided by a thin pad and a deformable headrest or a thicker cushion on a more

rigid headrest. For a rigid headrest, the provisions of Section 5.13.10

should be applied, and at least 1.5 in. of cushion is desirable. If the

space limitations of the application prohibit this thickress, thf cuu,hion

should be at least 1 in. thick for compliance with MIL-S-58095.

69
5.7 DES1M^'N PRINJCIPLES FO)R PFUjONMEL RESTRAINT SYSTEMI'

5.7.. ý;ent

~t~oeo'~nes.ses fi-wi perscnneI should provide the restraint neces-ýr,' tc

prevent. iT!juri--es to all aircraft occupants in crash conditions approac-h-ircj

thtc upper limit!; 2f SUrvivabil ity. Appropriate strength analysis and tests

a:; i'4:!cribed in Section 5 9 should 5(- conducted to. ersure that a restrair~t

systell is accepta-ble. Qualities that a harness ,,hould pos'sess ale listed

if, It should be comfortable and light in weight.

0 It sh~ouid be e.isy for the occupant to put on and take off even in

the dark.

0 It sl~ould contain a single point release sysftem, easy to operate

with ore (either) hand since a debilitated person might nave

difficulty in releasing miore2 than one buckle with a specific hand.

Also, it should be protected from inadvertent release, e.g., caused

by 11h-ir buckle buing struck by the cyclic control or by inertial

loafiing.

* 'T should provide personnel with freedom of movement to oper3te the

aircrYaft con rols. This requirement necessitates the use of an

iricr Li c ree(,l ill conljUn.lt ion With the sloul dnr harness.

Il It should providi! sufficient restraint in all directionsý to prevent

injury due to dLCeleratlive forces in a survivable crash.

0 1 hc w' bb rin should pro..'idu a max imnum area, coi,s istcni ' with we yht

and( curiifurt , for forceU di stri but ion in the upper torso and pelvic,

re.y lots and should be of low eliongation unoe)lo1 ad to minimize

dynamici overshoo0t.

6.7.2 lye f yj;

5i.7.2.1 Ajrcru .y~em~s. illir exi sti ng mil Itary l ap tb-l t and shoul der liar11 LU coif i j ta t oi 01W i t 13L C r. tC 1- t if.dwr st-a as1 4 s hown in. F i (jiire ?23 s h oulId

he u',Led by U. S. Alliý my ilots. llic cot C ~ifiJUration) shown in I igurle 24 is preI lred: ecv it provicu ides 1mroved 1lateral restr'airut due to theý add it ion )f

thenui vvlv~ted shoulder strap%. ljhiý systetil esujl ted flrom the_ i live st i(j i on

~:i I t!di ii Ref (.tur. ree 4 4. Dueala 11, f the har-dwile itl tll'iný sy:stemsr arv It '2-

~e il i Sc_.t.ioni 7 .5 of Volume IV.

5.7-?.2 Yr.7 CYo.!'i.(.ndc-'atioris ill the :,vlcc:t ion of a troopl or pas1

si rojf- %Cat' ve.li~r lt sy'teIi a di I letcilit. froml tlio~u for all arl'(Acw syseiii.

I iI- 1t of all , t I I se at mlay fI.h(- f IVIMrd , s i!-a d':~ J, 0) aftl twI d . Seci -Lldly, the(

lcuýtIi d int sy,t, I-i mru,t L I .a pahllIe uof bLilty UAtal'cd oiid remove:d ilicly diii

op r'rl lioIi.Jl 1:11tlV -oltirt"iltl. Fiy 11.1(01 1oi)rrtihre ) vatyting tyi)"'s arid qiidtlitlit 1

of uquiiijriiii , Al so, whuri,.,:. a ) i O lot Proidtl/ ul'cs thul rt:!4irliit.l :,y'tetil ill

lis Jaltdft t ',i su I querit y th-1it Isu krf.co;i: a Iliittcr of 11,411, U~ollp. anld

nasiniursayi oft Let niftminl iii whil- thir. systert. ille vff U t- oif thii 1 l1 k oif

f.,1ruM 1 LIity Wi,IJld Ip utrhuly biecotric 11101UL.r utoliJ'lM. cd iin a '.ombat %lI uit lon Ownr

7 (
ITEM IDENTIrY

1. BUCKLE ASSEMBLY

A. SINGLE-POINT

RELEASE BUCKLE

B. TIEDOWN STRAP 3A

C.. ALJUS rOR

2. LAP BELT ASSEMBLY

A. ýAP BELT

B. ADJUSTOR

3. SHOuLDER HARNESS

ASSEMBLY

A. INERTIAL REEL D

B. INERTIAL REEL SRA

C. LOWER SHOULDER 3

STRAP 25

0. ADJUSTOR 2

FIGURE 23. BASIC AIRCREW RLSTRAINT SYSTIEM. (REFERENCL I-)

thIJ v-i,,k ii fvr)*ved in not u~.iwn thIw rustralnrL systcmi I'c.uri'rs uvI: h uqor.

11Iv),efufT';, iIalVdWarI: slioIlG d I LlIKcUllijil icaLud an~d, if Vo,,iVe UB(Amb1c 1,t1llc

No( syýLum5i tildt vu.,ullIud fromr the lilvesýtiq.,i. ll f2w~itud Iin Pehl 11c 4!. ai ':

shownil Ii igiiru ?!). 111" 1YpI' H tr oup vcetta3irt Sy',ICrrI il jircLfQerre anid LUll

~1~ of a two strapi shiuuldei Ildiw1i2, arid a laj b)121 a-isliihly. *the, two~

01miUIlJ dus I all" a dL: att Lhicud tu tVI') sItiI(j I ( hi ll 1. I1 1 1 ,ýc .. .I lI (-y i~xic-ld tIm

waird arnd domi ovc-r the uc~cupaiirI 111,11f. tol-,li itid w.11:l.llI( intoihi i:

I('f I. dliii ty ihiit b)121 , with adjjiil"Iýs th 111 dl12tlý l~lI loyjl thill at

71
ITEMk IDENTI7TY/4

1.BUCKLE ASSEMBLY

A. SINGLE-POINT RELEASE BUCKLE

B. TIEDOWN STRAP

C. TIEDOWN ANCHOR 4

2. LAP BELT ASSEMBLY

A. LAP BELT 3

B RETRACTOR

3. SHOULDER HARNESS COLLARA

ASSEMBLY

A. PAD

B. ROLLER FITTING 3

C. ADJUSTER

D. LOWER SHOULDER STRAP

4. INERTIA REEL ASSEMBLY _1

A. REFLECTED STRAP

6. ANCHOR2AI

C. INERTIA REEL

FIGURE 24. AlkCREW R[STRAINT SYSTEM, INCLUDINGi

REFLECTED SHOULDER SWPAPS.

d if f urý I rumt the lypt: I I re~traiiltt hy iaV ingla si Iitly S hIculdoI- str a I t IIi 11

paI %vsC d i a yfol aI. ly da U 0ýs t IlIv! (LLI pfInt t up U p (.I. tuUV!U. I Ur - (kt f 'J(' 1l sej Zit

ji I s II )ulId 1 '5 0V('I I tIt d~ Uv (huu10iu U 5 1.t tU t IlIl: I Uw:, t I thu di I (. v-dI t . ]If 11i0

'IyIIC, I syACIII I U!,c( ill i titit(t.v a fL'Wa? d or aft. fI I 119( ýuat, tiiw ( Idk-'Jolldl

O~ilo I do. s vI tipt -.shtl U i'- a uVI: tltt: (ut~tiJurt shou itii to vc-:.Lrjhi tiW ULLU

halIt I 1Um~ prutrudi1119 uutsidft: thu dHil dl L dut iriy 1ot urld lujd ilic.

fur LI ! 01 Litluf ',antido U I/WiIldtlw yUI! ?u h~ 111 i i to (1.41 sy.,tuiv. " hoiw

t!VVl , titt.y ;IIuIIt a]i IJ ow i1_111, (A 111111:t,1LI to I~ll)VC. out of I. ho SI!dl t u 't(.)f uflill

I .i1td In I i ll UI 1111) uji.II CU dl:d5S lIt: s y %t~u ( il 5 I'uu 11( 1( 11 tLt d il tI ti u t. L~ I 1.t oI ti
~~ 2) 1 ITEM IDENTITY..,

1. INERTIA REEL (OPTIONAL.)

2. SHOULDEn STRAP

3. LAP BELT ANCHOR

2 4. BUCKLE WITH SHOULDER 2 STRAP CONNECTION

5 LAP BELT

6. ADJUSTER/FITTmO J

33

4

6 6

5 4

TYPE I lYPE I

FIGURE 25. AIRCRAFT TROOP/PASSENGER RESTRAINT SYSTEMS.

seat the instant he returns to the seat and provide adequate restraint. during

a c.rash . The systemi should mai~intain0 tile 1lap belt buckle in the proper rel ati on-1hip to thle guniner, pr event inrg the shoul der straps from pulling it up or

the 1 ap belt f'rom IMuIIing-1 i t s i dQviayS . Such a systemr has been descri bed in

RuI ererice 46 anid is ShOW11 inr I i gure 26. It con i sts of a lap belt with

irioriy'1 ta- C-Zr ! :;ri o4 i f-" Lrn p It W(i~ !,,rc v in

* ~ar inrverted Y arr-anyrnrent, to a sii oqi itirert ia reel Strap. I1he lapi bellI. with

thighk si rap atid'lirrr11errt is ea:sy to put, oi arnd pirevents the 1,q )0 bet, f rom rid

ilig up11 dUI-ifj 01)operati on 01' tiý grin. lh1w lap belt, is ~l u(gijed inito tire twoý

seat pan ill(- rt ia reel s whi-i the L1rewmrerrber is, to 1,eu seated of- is standinug in fronr. of tho seat.. Iliw. siou I del 1 ij vrins s arid 1 aj bel1t w I th tii igl st10rap% lildy

ssevve asý ai Imnrkey 1havn(iiýS" Wii(!t tiQ recrwmremrber di-sConiriects tire two lap beclt

Ill iyj lit I itt irvy, lu ton liermit. id ree-ls. lIire vresul toot. coot i~lor;t ion p~I" rlliit5

tie 00I Irri~rrii rodei)trAo V IIIl( ye-IJV tiavi I withfin tin c iii Vl wi ~still kifeinr -oil

rleActd to timý &ioulder1 irarrrrs lrr'r I. a reel, lirerebCIy ros~tra iii I ri tire Cre(-W

r Irr~ worn fol irig out. of tit(. a i i-taft

5.7.2.4 1!'fIg 1Ab1 a_ c Sys t V1 , .i All o~lorrat jt. at Iy I n fIa ;t a1)l-, body t-1rd Ifea d re,-

s t tjia jirt, ~yt y. (-r ( I I3Ai 116) I c) I If( i(.Ojnt (I- -I-(wrrirIerr a. h celler Jr nitly dt.velo~(-jrr drid

by:,~j i 1.1i. rN'iv.i A ~ir [lhvi opmrrirt (.ert-Q arid lie Avi-dt ira App1 ktd lec~mrrl

irrj~y rIm tctui .11'!. A%'.1 if, lnrtr od if] I 1gmif ?/, 1in K s~y-Alrr i rrtV;nles " ilrr !,-(ieasJ

(r.1 " I1 1 1i c t.f:( t, I (i w irc)1.ca'j ", -I it 1, OY o v I ii tI-r I L L I 1 ri r piA r.ira. i oirr i rit I. that II Ui k.c2 Lii':

o (t1 1 -rjinl L Im .ki. Ir irI IIs 1 ,11 , t 'ir -i t -by i I. I (i I I d' (yrII, iarrIa ove ~I. ImIIo I t ri d II od du I o I r Ig 

%tIt .11 1(), i lli.)bi oll till, w'.-.1 et wlirer tiul- [tt 1 0ted I Lsti .1lilt Is " crri e-llptC ,d riot log'
2

54

2. SrOULDZR SiTRAP L 3. SHOULDER STP.AP ADJUSTER

4 ATTACHMENT RELEASE BUCKLE 0 0. LAP 13ELT

G. LAP BEL.T IMERTIA REEL

7- -HIGH STRAP~S

a. THIGH STRAP ADJUSTERL

V. LAP BELT PLUG-IN FITTING

FIGURE 26. GUJNNER RESTRAINT SYSTEM. (REDRAWN

FROM REFERENCE 46)

the crash. ihe concentration of strap loads on the2 body is reduced be~cause

of the increased be~aring surfoice p~rovided by theý inflated rcst~raint, arid buLb

1k(aJ rotation~ and the p055 Uibii ty oif w i p1ash i nduceud trauma are al.so

Al thonugh ImIOr omL0()',X dfi(I Co.;tl y thwil convent iuntal re Ar-ai ut s)'l teiiis, SU i a

sy .~tumi way bu Juo s.if I ed becau so of it% potin t ial for 1impr oved occ(upatit pv )

1,U0I. i On - OcVel 0pMI0rt Of thL sy ~tewil anid results of test I frPJ are documenltec In

Icf urun(ces 47 dnd 45.

6.7 .3.1 1 Lit oIbv hbu- tesre,01' coIIf Grt IItU',t 1i(0t tIC CLllijit 1)111I std by

(.C ui v Vl'dl I I IL(1ttfttI I 1if-ttlit :111, 1. dii 1 10iif V1 I conts'Jur.'At Ileti fu0"r u :t) IA dint.

Int tesesi.- the ;jbIun(qIL of ri'jid hwtdwdr(. e 1uicitd uvt.,i bony ittv lion, of the

74
RESTRAINT HARNESS/BLADDER

I ~INFLATORS~

lit

_~5PON ROTARY BUCKLE - 1~~

'~7~I~ ~ (MIL.S.58O95) /

CR1HSENSOR4

Iii,

'.,.. * >'- AIRCRAFT

NORMAL D.C. POWER INFLATLD

FT GIJRE 27. INF-LATA\B[E BODY AND HEWAD RESTRAINT.

(REFERENCE 47)

tor~(). Al ,u , har'Jwarc ti.cih kthdt aY'(. toJ W idi2 oV I al '7 ,Qo ar'j 110!. Cu-rt -

fiqlIt (2d Officioritly tu fit t-[o:. dcJircU ltdt ioh (Oil the body, LQUld bQ Urow()iii

f ortabil (2. wchh i rit thati. ' too' wide or too stiff could ai ,o (airI:I di-,'xurf ov

th riu(jh I c i's i nj (if t ho whbiw) i yr V 'Ljrat Iion due' tu Ir udutoiJd vkut ii Iit onti

53.7.3.2 FjI11ryý1 y 1t'U; 13Vquj~r1rrcIlvn. A snl"uIdir atiy Ia I': it U(,o~bi

113 1.i IjI sI IuMId I "t I v c a s iifl c( poi lit (if tIv (A t t h, c.a(ll L$: Or-I, a tod by I- l I1(21

handr so' tha t d0:h i I i totu If.O'T cdii . (O 10.1yf It t Vl(i-,:, f~ lI Q thl'i2 I'r t

st IIi aLII t'':cau',i -o 0f UIC ddtly' I of 0 ioJt tt .ii I wi., w- .,it:kingcj. flowovV'l , v-.niI

I i oiI , U(: L Iele 61. ;Vii Olo d i Ily, o' C0oIIi..i~t W~ tb 1tl( lii t O Ml.Ip~I)II Of jil 'I a A t(, i . t 1,0 1h ( .

01110d ['(It ridI~IV't toth'1-y ojWIr the Ilu(kh(', ;aid I thC liter!l-. itill i rv ll ,' ". lb.1)

lull'.) of Lf c'. I'If tiot- , tIhe leti.Jo Thould 0-H's~hi with"I it-' wuoilp:it.

1lrqil itivii lt(d IIIo lit I w A U liItI Sy-Al ,'11 d li if ')(IX * i IVrI(. ~ dl ~ '0 V( 1 0 S VIt Vv

7'
In restraint systems other than the Type I of Figure 25, if a lift latch or

similar type buckle is used, the restraint system design should ensure that

the latch lifts from left to right on all installations. This will reduce the

possibility of reverse installations and the resulting confusion.

The release buckle should either have the capability to withstand the bending

moments associated with deflections and motions during loading, or it should

contain features that allow the fittings to align themselves with the loads,

thereby reducing or eliminating the moments. If belt loading direction is

such as to cause the strap to bunch up in the end of a slot, failure can

occur through initiation of edge tear. The fitting and motion angles illustrated in Figure 28 are recommended.

'M/ S $ OULDER HARNESS

FITTING

LAP BjELT FITTING2/

7N (MOTION

RANGE, TYPICAL)

&.iJ "SINGLG -POINI ~ I{.L LAL LZUCKLL

FIGURE 28. BUCKL.U FITTING AIACHIIMNI AND IMOTION ANGLES.

If th' iir:t1ujr ity uo" the attaLhmjiwr:nt of .lu,: t itl in. within to I)Utkl(: (-l l I}• L

,.AjlI)qlhli',lPd by 'utitionl, thfii ro l~llati (J should be Lol ,:tltl'Iy LQ1 ifited

I I imli .ail1ij f il.ti i.?. I' olld i in ilh. 1Al I la o uf tli lw( kl d un ri, i 1 dinfj

II;'y ,A(YvO ' to I-n diJ iflu it. n l1 i(Itw ly!,l y.u ,' I. LX-l,'. I 1 ,',. ýIlovll th,l

it I , I,,ltf' to df .,yll U. adtla( hi,:iit of 1., i ill i Uj withi th: bhv.kl c Io he

iti.. n-,Il 1v- tv ti, ui h tt n t( y oi:ly..,,, the. t i l. l. In(j at. ',
rotation. F ir example, a round pin in a round hole would be preferable to a

flat-faced doc: which must seat on a flat face of a slot. In the lat-ter case,

a small amount of rotation can cause point loading of a corner oi the dog

against one end of the slot. The noint loading can easily increase the

stress applied at the contact point to -its ultimate bearing strength, which

results in deformation and the formation of a sloped surface which can act to

cam open the attachment m2chanism.

Further, the release mechanism (buckle) should be protected against accidentai opening. Neither decelterative loading of components nor contact with

aircraft controls, such as helicopter cyclic sticks, should open tha device.

F.equired cockpit dimensions should be reviewecd. It appears that the occupint

can be placed too close to the cyclic control in helicopters and that a ful ly retracted cyciic head can contact the buckle. The buckle release nl2chanism

should be protected apainst inadvertent release ei4ther during operation or -in

a crash. It should be emphasized that, if contact between the cyclic. control

and the buckle is possible in an operational mode, a considerable overlap can

exist during crash loading when the restraint syst-em may be deformed forward

several 'Inches.

5.7..3.3 Lao Belt Anchorage. The a Ctual anchorage point for the lap belt

may be located either, on the seat bucket or on the= basic aircraft structure,

although it is usually desirable to locate it on the sea. If the anichorage

is located on the basic aircraft structure, consideration should be given to

the miovement of the seat when load-'limiting mneans are uSed so that the lap belt restraint remains effecti'jP rpnardl1⁄2_,s of ýea poif-inn Lon~litld i nal

load limiting of the seat serves little pur'pose if Lhe lap bel t i1s attached

to the basic structure; therefore, thc belt should be attached to the seat

bucket itself.

The lap belt should be anchored to provide! optiniuin restraint. for the lower

torso when subjected to eyeballs-out (-Gx) forces. One of the anctiorage variables which has an inf'luencce oni iestrairit optimizaticn is tho location of

the l ap belt anchorage in the fore-and -aft di rectiori . ito imn'portarit charac.-

tori stic is the angle in it vertical force-and aft p1 arie butweenui a projection of the lap belt ceritc'l ine and the buttock refere:nce line, or pl ane. 1 h i

angle defines the geomnetri cal rel ati onshi p betwoeenr the 1 orni tud inal and( vertic-al coitiponents of the be] t load. lk sninai anýjl o prov ides at' ef I - ci en!. path~l for !-uppoirtiny i onqi tudi nal loads whilQ a large. angl:ý prov ides an c-fficieiit

sytoui for supporti rig verlt ic0 loaids. Thus, for supporting large I onwardecto(ed loads, a smal l anq Fe, woold Ic hi. desirable, blt. for reacti ng the 1large vertical loads irnipos~d on thec la[) belt Iny the loJadedC Shouldek-r- hiarneIss 1. lat ij

Sdrall th'.. vali abl s i nclluiintr the_ 1tc(nAlr.ny for theý o(rccpanil. to sutbniar ine

unde~r t he 1lap bul t

Cuinfuf tt I,, dnuth-r' curorci rinI lap; Felt atilc~or lod alIott. A p1 lut. rmi,J,. t .

al( Ii. o (v t iih iril. , o al ;otlU e du or ar .Ip'riill" . 11 i 1i,

I'l beil it alit. hor I : toot I itt f ij-Wd) d t hP I 1 u: a ) I, Fme I I W ~ i s tVr I ll 11 t )01. S

Itli ghjIt (Jo I wa I d 'J i Ht .- ( (:", t A. h ctM(1 1.11 1.1, i1I' .Jli' J ') It l jI ,.V I t, '11 1.111. 1, 1. (,

fet (- wlIli -;(:t Lical I1, tioo oitI. 1u 1q p 0~ u Iit rt , e-fov to Ilu., if olou 11i
lap belt anchorage so that it provides optimum restraint while not interfering with the pilot's operational tasks. A more forward location of the

anchor point does not reduce the comfort of passengers, since they do not

perform such tasks.

In order to satisfy comfort and crash safety requirements, the vertical angle

between the lap belt centerline and the buttock reference line as installed

on the 50th-percentile occupant should not be less than 45 degrees and should

not exceed 55 degrees, as shown in Figure 29. Further, it is desirable to locate the anchor point at or below the buttock reference line to maximize comfort anI performance. If the anchor point must be located above the buttock

reference line, as on most armored seats. the anchor point should be positioned to ensure that the belt angle lies within the desired 45- to 55-degree

range. For a system having a lap belt tiedown strap to counteract the upward

force of the shoulder harness (e.g., in pilot seats), i.h2 lap belt anchurs

should be positioned so that the cente,'line of the lap belt passes through

the seat reference pcint. If the restraint system does not have a tiedown

strap (e.g., in passenger seats), the lap belt anchor should be positioned so

that the belt centerline passes through the buttock reference line 1.5 to

2.0 in. forward of the seat reference point. This position provides sufficient vertical load to counteract the upward force of the shoulder straps.

K (S

StAT.

TAN-EMN LINI,

1<1 51-

p9( itj "/ ~ (Ii I()(.;K HL 'I I I fJ., 1. INI,

I 2 AP 1i / (J IUi

t ?J. ludP, LI, AF" lLi~l AF, O 7 dA L G LH t Y ( !L L L
For anchors that: do. net'I fall on the buttock roference line, the angle between

the lap belt centerline anid the ý,uttuck reference line should be 45 dlegrees

for systems with tiodown. straps and 55 degrees for those without.

For spats that imit late-al motion of the occupant with structure, such as

in armored seats, the anchorage point and hardware should possess sufficient

flexibility and strength tc sustain design belt loads when the belt is deflected laterally toward the center of the seat through an angle of up to

60 degrees fron: a vertical po'sit-ion. lhe side motion of fittings on other

seats should also be capable of supporting design loads with the lap belt

deflected laterally away from the center of the seat through an angle up to

45 degrees from the vertical . These recommendations are made to ensure that

late~ral loading on the torso will riot result in lap belt anchorage failure.

5.7.3.4 Shoulder Ha~rness Anchorage. The shoulder harness or inertia reel

anchorage can be located either on the seat back structurel or on the basic

aircraft structure, although it is usually more desirable ho locate it on the

seat. In placing the inertia reel, strap routing and possible reel interference with structure during seat adjustment or enrgy-absorbing stroke of the,

seat must be considered. Locating the anchorage on the basic aircraft struct~ure may be the only practical approach for improving crash resistance in

light aircraft, particularly in retrofit applications. it will relieve a

large portion of the overturning moment applied to the seat under longitudinal loading. However, due consideration must be given to the effect of seat

bucket movement in load-limited seats. Vertical movement of the seat. can bao

accommodated by placing the inertia reel a sufficient distance aft of the

seat back shoulder strap guide so that scat vertical movement will change the

horizontal position and the angle of the straps very little.

Shoulder straps should pass over the shoulders in a plane perpendicular to

the back tangent line or at arty upward (from shoulders to pull-off point)

angle not to exceed 30 degrees, as illustrated in thae upper-left sketch in

Figure 30. A shoulder harness pull-off point should he at least 26.5 in.

above the buttock refrened 1 ine; honwever, this dimenbion should not be

increased, because then the harness would ot provide adeqibate restraint for

the whortrut occuparit.

The shoul der harness anchorage or guide at the top of the seat back should

permit no more than 05 i. lateral moveiert (slut nro morin than si ian. wider,

than strap) to ensure that the seal occupant is properly restrained laterally. ohdn guide should provide smooth transitions to the 5lot. he transie

tion contour shiuld bsa of a 10e VadiUSo t an 0.25 in. and ots should extend

acomplmdtely aroupnd t pephery t Of tia reSlot sufiientist to edane War on the

strap and reduce tlder possiia lily of thbbing failure due to contactw with Sharp

odgeis under hpit l1noad itri. Also. tle guido e thita th strap load.s should b Suldiertrapy stiff to limit del lectioh under load. lane deflectiar ta

produce edac t angeldrigj arndae o atnprmatured (f f sa ou ero O ti wlebbilng.

5.7-3.5 1 qp . U.1Jil!JQw .§Cp)Anc-hqraqe. A 1 l) 11121t ti('doWri strap is

neonti t i3or eed forward degr ey ( ) llustrnibe -. d 1 he tiewnup traleit asktohra in

poFinet. s d 1 ho u:lderhard oif I shuat pIoi ctili at bei a point 14 to 26 in.

forwao d of lie seat hre n I oe lshortur seat dlli"e, the anior rushudnt be pCed

ith fil ortewr o ud pan t. j flL

79
Correct

Backtanent ineTorso carries small

Maximum vertical load

sShoulder

straprp oa

.. ... ..... ....

a n .. 4.'.. ... . . . l 4 h ul e

.. ... ..s r p l a ... 4 . . . .. . .

A (Foce digram

273~0.5 0

Sea..refer.nc.

point

. . . .. ... . .

o . .. ... .... ..

..... .. .... .. 0 5Cin: 

27.0

. .. .. ......

.. . . . .. . . . .. . . . . . . . .

. ... .. .....

................... t

.I.... ....... ' L ~ i J d

.3 .... ~ ...... ....... D L .N ....... ~......I G 

.. ....
5.7.3.6 Adjustment Hardware. Adjusters should carry the full design

load of their restraint system subassembly without slipping, crushing, or

cutting the webbing. In extremely highly loaded applications, this may require that the strap be double-reeved in a manner that allows the adjuster to

carry only one half of the strap assembly load. The force required to adjust

the length of webbing should not exceed 30 lb in accordance with existing

military requirements for harnesses. Insofar as possible, all adjustments

should be easily made with one (either) hand. Adjustment motions should be

toward the single-point release buckle to tighten and away from the buckle to

slacken the belts.

5.7.3.7 Location of Adjustment and Release Hardware. Adjusters should not

be located directly over hard points of the skeletal structure, such as the

iliac crests of the pelvis or the collarbones. The lap belt adjuster should

be located either at the center of the belt near the release buckle or at the

side of the hips below the iliac crests, preferably the latter. The shoulder

strap adjusters should be located as low on the chest as possible.

5.7.3.8 Webbing Width and Thickness Requirements. Webbing requirements

are discussed in detail in Section 5.7.4.

5.7.3.9 Hardware Materials. All materials used for the attachment of webbing (release buckles, anchorages, and length adjusters) should be ductile

enough to deforri locally, particularly at stress concentration points. A

minimum elongation value of 10 percent (as determined by standard tensile

test specimens) is recommended for all metal harness-fitting materials.

There are obviously some components that, for operational purposes, rely on

hardness. These components should be designed to perform their necessary

function but be made from materials as nearly as possible immune to brittle

failures.

5.7.3.10 Structural Connections

5.7.3.10.1 Bolted Connections. The safety margins for shear and tensile

bolts in restraint systems should be 5 and 10 percent, respectively. Also,

bolts less than 0.25 in. in diameter should not be used in tensile applications. Wherever possibie the bolLs should be designed for shear rather than

tension. because of the vibration environment in which seats operate, all

fasteners that affect the structural integrity should be self-locking or

lock-wired.

5.7.3.10.2 Riveted Connections. lhe tpjnidelinres presented in MIL-IIDBK-5

(Reference 43) are reowmciirded for rcstraint system hardware designr.

5.7.3.10.3 Welded Connections. Ac(ceptahile weldinq processes are discussed

in Mil itdry Splcif iuatiion M1I -W 860N , -60873, -45205, arld -8b]I; hUwever,

strict iri",lo:ctiun procedurus, shuUld be u',ed to ersure that zfll welded jointr

are ol adequate quality. (Other provi sion•, prreuerted in Section 5.4.3 al u .'JI~lly. )

5.7.3.10.4 Pl astic Strength Analysis. Pla itiu analy-,is w 'thuds 5houUld

1w u'ed for slirru'j h d(l.livmiiiOi in wl'i ,ever appl •bhl, ill order to )l1air

max imuu si r ji Iitt,:ngll, haidw •i ,at tbh 1 Iw,.t po',sw hIe w,-ight.
5.7.4 Webbing and Attachments

5.7.4.1 Properties. The main advantage of a single-strength harness (only

one restraint harness in the inventory) would be the assurance that harnesses

could be interchanged between load-limited seats and nonload-limited seats

without fear that an understrength harness might be installed on a nonloadlimited seat. On this premise, the design strength of all forward-facing and

side-facing restraint harnesses should be equal. The design loads for the

various harness components attached to the seat are listed in Tabic 9. The

elongation of all webbing used in the harness must be minimized to decrease

overshoot. Table 9 shows that the shoulder harness elongation is restricted

to 8 percent, while the lap belt is restricted to 7 percent when stretched to

a load of 4000 lb. Restraint systems for the new generation of Army helicopters use a low-elongation polyester webbing, other characteristics of which

are also listed in Table 9.

TABLE 9. OCCUPANT RESTRAINT HARNESS REQUIREMENTS (MIL-S-58095A(AV))

Harness Webbing

Min irnum Harness AMsembly

Tensile M i n imum

Nominai Br edk i r: r.. iu U U ti m. ti

Width Thickness Strerngth Elongation St rermjt h

Component -n ._ ( i .,)__ (b y) - (1h)

Inertia reel leid-in 1.15 0.055-0.075 8,000 8 @ 4,000 lh 5,000

Shouioer harness 2.00 0.045-0.0G5 6,000 8 @ 4,000 lb 5,000

tap belt 2.00 - 2.?5 0.045-0.0G5 J.000 7 ( 4,000 lb 4,000

Lap belt tieudowr 75 - 2.00 0.(4S-0.0jS ,OUC0 10 I " 3,000 1i 3,000

(1) u drO terinirne elongatih or o , d ,ilul ol , ult ilu.it(: struryi H, thf- ,IUu, lJ•':r haT: -cý,"

a,/,lXjly 'iand the in t ' . 1et: L , : h3: 1 ,ha l l eot: ed t•r I' hur mu n i. rj igut tin'. t, if W .

t hu I 11CI t . r-, -1 I r i , ,b j 131lkd I 1,0 1t I .urdI j Irtt.J, hl'd to a 5uItulto)1C 5tti t rI y i

t u , liU. : t') :,/oluld,:i ,i fln :, end f itt Ifl,, , I Fhl I l[, ',lwjj,]t:ui ,tr I h: ,Il.i: Ir:

-,I'l th . luM1' ,f attic ,cd tI ,j iov.)l, f t Ixtul U: l 1' i f h1 1 i 1 i) : iu'ui'.J Id to

1i . 5 ,t/i p,:rLu,10 IlI, OW.C.u,. ,iiu . If:: tI' t ., / 1 1" ,•,() L d ,I, 'k .. { 1,-1 iJ I II .

i¶ bri 4 7/./.3 of Mi *S rbtu'jSA, (t,V) . r,,'J fill: cluqu.p t uou "I"),, l t iJ'i1-[ ' )m ,,d I t1.1:

I: it ý1u)1) 1-j l hjI I Ii , LIU',b V 0V f tI e Q , 1 1 1U j , -i,' l ij ! (1 to I I, ,L

(7) 1 ' . l, ' d.t/ . tc .1.lt uI lIhUll, ",Iu t o 'tII:b I lh. l tnw" I •I, Ill j I I:i l l •l, f.l 

,: hluu t,i '1 000 ,hiJt V0 1i . I h /Ic,- 4 I ,',, t*,iuuulu rI 't.It, ./ i 5000 ur, Vd. , I br ,b: u

i (, I L o 1 , ; ii/it u 1".,t r i.h lt l-l, ( to u , iuulJ'r .)'.tI,|I t •h.1 "11, -, 1 ( ,;,' •lul- l•I u I ulh J'H,

o _, h 0tl, 4 ' 1 1Auýu ,1,
5.7.4.2 Width and Thickness Requirements. Minimum webbing width requirements are specified in Table 10. All webbing used for restraint harnesses

must be thick enough to ensure that the webbing does not fold or crease to

form a "rope" or present a thin sharp edge under high loading that will cause

damage to soft tissue. A minimum thickness of 0.045 in. is considered acceptable for the shoulder harness and lap bell straps, while 0.055 in. minimum is

specified for the inertia reel lead-in strap.

.0 TABLE 10. MINOIUH WEBBING WIDIH

RCQUIREMENrS

Minimumi Width

IWebbin IdentifyLn)

Lap belt 2.00*

Shoulder strap 2.00

Tiedowin strap 1.7,

*A greater width (up to '1 in.) or pad it.

desirable in the Lwitvr abdominal area.

5.7.4.3 Webbind Attachment Methqds

5.7.4.3.e Stitched Joints. The strenbth and rlsability of titchedses

must be entured by using the best known cord SineS and stitch pfol trnse or a

specified webbing type. The stitch paaternd and cord sizngs used ir: existinehiyh-strength military restraint webbines appear to provide satisfactory performance. The s lasic stitch pattern used in these harnesse . i a n. 11" con is

uration for single-lapped joirills. The 27-lb) strcnqth No. 3 nylon thread at 6

to 9'~.i t ches; pf~r inch is, recommendj(ed, as I 1srti il''ru 3! , furus

6 ~~on Mil_-W-25361 wObbings. The use of the 27-l1) Lord and an 80-p~rcerit effi

ciency refrults in a eriniaum strcnyth 0 130 la-/in. (s0 x ?7

x0.80 efficiency) for o >ingjle lapped jontr or 260 lb/in. fur a loopedt

j uint1.. I hus , the tot~al Stitch 1 eýngtlr iiucdcd 1-ar he tterrriined thy tic total

required load.

1t, has been shewn recuntly tha-t the .av iur thread Is riot cuonpiat ibl wi th thc

new low el onlat ion tl yes urwbb inq( ;~e Urre4) . [or thlvS'2 wi~bbi ny 1, a

smal 11 e r d i ame te r cord o)f i- us the iadvaritiayrs ofI re(do cid we 1)1 nirig f ie du GIUIJI&

arid to blY eued WiUth autOiatiL iMnu Wid tHi

aebincIdntty _bi. ne_.

Lap bet 2,03
-- 6 TO 9 STITCHES PER INCH.

FED- STD- 75 1, TYPE 30 1

METALNO. 3 NYLON THREAD PER

MEALDAR V-T-295E, TYPE II,

HARDWARE •CLASS A OR

EQUIVALENT

"0."12N. 0.25 IN.

MINIMUM SPACING

1.5 INIM ..... ONE STITCH

MINIMUM MINIMUM

FIGURE 31. STITCH PATTERN AND CORD SIZE.

The use of a 30-percent increase in the total stitch length required is recommended to offset the normal aging strength decrease as well as the possible

abrasion strength decrease. Covering the stitched joints with cloth to

provide wear protection for the cords is also recommended.

The size of the overlapped and stitched area should be minimized to reduce

weight, reduce the stiffened section of the webbing, and provide more room

between fittings for adjustment.

5.7.4.3.2 Webbing Wrap Radius. The wrap radius is the radius of the fitting over which the webbing is wrapped at buckles, anchorages, and adjusters,

as illustrated in Figure 32. The 0.062-in. minimum radius should be carried

around the ends of the slot as shown in Figure 32 to preclude edge cutting of

webbing if the webbing should be loaded against the slot end.

5.7.4.3.3 Hardware-to-Webbing Folds. A possible method of reducincj fitting width at anchorage, buckle, or adjuster fittings is to fold the webbing

as shown in Figure 33. This reduces the weight and size of attachment fittings; however, it can also cause premature wehbing failure because of the

Force applied by the top layer of webbing compressing the lower agiinst the

fitting slot edge. If this technique is to be used, tests to demonstrate

integrity are recommended. Also, for configurations that require two load

paths, such as lap belts, where an adjuster cannot hold the required 46OO-lb

load, the webbing is looped through a full-width slot which halves. the lo'd

in each strap. An adjuster is then included in one strap. Adjustment requires that the webbing be freely drawn through the fitting, a requirement

that folded webbing cannot meet.

84
A

RRF _4•, C=._WEBBING

WRAP RADIUS (0.062-IN. MINIMUM)

ALSO APPLIES IN CORNERS UP

TO 90" FROM TENSION AXIS

DETAIL A

FIGURE 32. WRAP RADIUS FOR WEBBING JOINTS.

W

FE

2

E-E

FIGURE 33. WEBBING FOLD AT METAL HARDWARE ATTAC.HMENT.

5.7.4.3.4 Surface Roughrcss of Fittings. A surface roughness of no more

than RMS-32 is recommended to mrcvent fraying of the webbing due to the frequency of movement over the metal.

5.7.4.4 Energy-Absorbinq Webbing. Energy-absorbing restraint system webbing has been considered for limiting loads on the occupant. However, prilnarily because of the increased potential for secondary impacts of occupants,

85
energy-absorbing webbing is not recommended for use in seating systems. The

limited space available in aircraft requires that the strike envelope be minimized. Therefore, the use of the lowest elongation available is specified.

5.7.5 Inertia Reels, Control. and Installation

The inertia reel should give the crewmember full freedom of movement during

normal operating conditions while automatically locking the shoulder harness

during an abrupt deceleration. The design requirements specified in MIL-R8236 (Reference 50) are compatible with the restraint harness requirements

listed in Table 9, and it is recommended that the use of this specification

be continued.

In addition to the MIL-R-8236-type reel, which has the function of preventing

further strap extension, there are power-haulback reels that rapidly retract

slack to apply a tensile load to the belt. Generally, these systems, some of

which use a basic MIL-R-8236 inertia reel, are powered by a gas generator and

must be manually actuated prior to impact. Automatic actuation by an acceleration sensor is not recommended, because human tolerance considerations limit

the haul-back velocity. By the time the crash could be sensed, there would

not be time to complete the haulback within tolerable accelerative limits.

It is recommended that the rate-of-extension type reel be used on all aircraft types to assure locking regardless of load direction.

The inertia reel may be anchored to thn snat, hack structure or to the basic

aircraft structure with the same reservations previously mentioned in Section 5.7.3.4. The shoulder straps mist be maintained within the acceptable

angle range as presented in Figure 30. IF an anchorage to basic structure is

used, consideration must be g~ven to the possible seat bucket motion so that

the shoulder strap remains effective during the energy-absorbing stroke. The

reel should be mounted and the webbitg routed so that the webbing does not

bear on the reel housing.

5.8 SEAT STRENGTH AND DEFORMATION DESIGN REQUIREMENTS

5.8.1 Recommended Occupant Weights for Seat Desiq.n

The 95th- and 5th-percentile occupant weights are recommended for the upper

and lower limits of occupant weights to be considered in seat design.

5.8.1.1 Crewseats. For some applications the design weight should be

based on the typical weight of the seat occupant, not the extremes. This

means that the aviator weight recommended for crewseat design should not

include combat gear. Typical male and female weights are presented in

Table 11.

5.8.1.2 Troop and Gunner Seats. The same percentile range of occupant

sizes should be considered for troop and gunner seat dcsigns. A greater var.

iation of clothing and equipment is used by troops than by aviators-; troop

seats should be desig, ,d to accommodate them. The 95th-percentil, occupant

should be considered heavily clothed and equipped, while the 5.th percentile

occupant should be considered lightly clothed and equipped. The typical

weights of male and female seated troops in aircraft are as shown in

Ta~le 12.

86
TABLE 11. TYPICAL AVIATOR WEIGHTS

95th- 50th- 5thPercentile Percentile Percentile

Weight Weight Weight

(b) (lb) (Ib)

Ilem Male Female Male Female Male Female

Aviator 211.7 164.3 170.5 131.4 133.4 102.8

Clothing 3.1 3.1 3.1

Helmet 3.4 3.4 3.4

coots 4.1 4.1 4.1

Total weight 222.3 174.9 181.1 14?.0 144.0 113.4

Vertical

effective

weight 175.2 137.2 142.3 111.0 112.6 88.1

TABLE 12. TROOP AND GUNNER WEIGHTS

95th- 50th- 5thPercentile Percentile Percentile

Weight Weight Weight

(ib) 0(b) (lb)

tem Male Femal Male Female Male

Troop/Gunner 201.9 164.3 156.3 131.4 1?6.3 102.8

Clothing

(less boot3) 3.0 3.0 3.0

Boots 4.0 4.0 4.0

Equipment 33.3 33.3 33.3

Total

weight 242.2 Z04.5 136.6 171.7 166.6 143.1

Vertical

effective

weiyht

clothed 163.9 133.6 127.4 107.5 103.4 U4.G

Vertical

effective

weight

equipped 197.2 1N7.1 160.7 140.11 136.7 117.9

87
5.8.2 Strennth and Deformation

5.8,2.1 Forward l.oads. For a load-limited system, a minimum displacement

must be achieved if the system is to remain in place during a given decelerative pulse. Actually, all systems are load limited, although not necessarily through original intent. The inherent load-deflection curve for any system imposes a definite limit on the system's ability to resist impulsive

loading. The objective of intentionally load-limited seat systems is to make

the best use of the space available for relative displacement of the seat and

occupant with respect to the airframe, while maintaining loads on the occupant consistent with the type of restraint system used and the occupant's

capacity to survive the loads imposed.

Design curves for the forward direction are presented in Figure 34, where it

is estimated that the requirements are not consp'- at,,,e for the input pulses

selected for design purposes. These are a 30-G peak triangular pulse of

50-ft/sec velocity change in the cockpit and a 24-G peak with 50-ft/sec

velocity change in the cabin area.

The static loads that the seat must withstand are obtained by multiplying the

load factors (G) shown in Figure 34 by the sum of the total weight of the

95th-percentile crewmember or passenger plus the weight of the seat and any

armor or equipment attached to or carried in the seat. For crewseats, the

weight of combat gear is not included (see Section 5.8.1.1).

Longi,.udinal displacement of approximately 6 in. for cockpit seats and 12 in.

for cabin seats measured at the seat reference point (the seat reference

point may be projecteJ to the outsidp of the seat pan for mcasu'eCmernt convoni

ence) is the practical limit, for seats in existing Army aircraft. Since

there is typically more room available in cabins than in cockpits, the advantages of longer energy-absorbing strokes can usually be achieved. Longer

strokes permit the absorption of equivalent energy at lower loads and thus

can serve to reduce seat weight and increase the level of protection offered

over a wider occupant weight range.

In viewing Figure 34, it can be seen that for cabin seats 12 in. of stroke enables the minimum limit load to be reduced to 15 G, whereas for cockpit seats

a 20-G minimum limit load is required with only 6 in. of stroke. The 15-G

and 20-G minimum limit loads fix the G levels of the base curves for the

cabin and cockpit seat, respectively. lhe available stroke will be unique

for each specific aircraft, and the energy-absorbing mechanisms in the seats

should be compatible with the available stroke distances, If forward or

sideward motion threatens to limit the effectiveness of the vertical energy

attenuating system or increase the possibility of severe injury caused by

secondary impact of the occupant with items in the aircraft, then energyabsorbing stroke in directions other, than vertical should not be used. The

6 in. and 12 in. allowed by the curves of Figure 34 should be viewed as maximum distances which are subject to limitations of available space in each

specific aircraft and location in the aircraft.

The initial slope o{ theŽ cockpit seat base curve to 1.0 in. of deflection

allows for elastic dcformation consistent with a relatively rigid crewseat

while the lighter weigiht and more flexible troop/gunner seat requires a

lesser slope. The 30-G and 35-G upper cutoffs reflect consideration of human

tolerance limits, load variations between cockpit and cabin locations, and

practical limitations of seat weight and evcessive airframe loading.

88
40 ~ACCEPTABLE FAILURE AREA4

0 (EXAMPLE FOR

______D FORWARD-PACING SEAT)

0l 30

'40 SAPL

C~~~ 25__L ~ ACCe, -

BASE CURVE

0 CODE FOR SAMPLE CURVES

19 1 0 SEAT FAILURE LOAD

_UACCEPTABLE PERFORMANCE

10 c'-.----.- INEFFICIENT PERFORMANCE

~ COCKPIT SEATi

0 LL

0 0.25~ 1 '43 4 5 6

35 -ACCEPTABLE FAILURE AREAI

0 REJECTION AREA

CJ 20 171.5

00

4v

6 CABIN SEATS

0U,5 2 4 6 8 10 12

TOTAL CONTROLLED DEFORMATION WX) IN.

MEASURED AT SEAT RE!FERENCE POINT

FIGURE 34. SDI)' FORWARD LOAD AND DEFLECTION REQUIREMENTS FOR ALL

TYPES OF ARMY AIRCRAFT (FORWARD DESIGN PULSE).

89
5.8.2.2 Uie of Design Curves. To be acceptable, a seat design must have

a chavacteristic load-deflection curve that rises to the left and above the

base curves of Figure 34 and extends into the region beyond the upper curve.

ThiF requirement also applies to the lateral strength and deformation requirements discus;ea in Section 5.8.2.6. In Figure 34, curves A, C, and E are

acceptable curves, but curve B is unacceptable because it does not reach the

requited ultimate strength. Curve D rcveals inefficient use of seat deflection by intruding into the base area. The seat is deflecting at too low a

load, thus absorbing less energy than desirable.

5.8.2.3 Aftward Loads. Large aftward loads seldom occur in fixed-wing aircraft accidents but may occur in rotary-wing accidents. A capability to withstand 12 G is recommended for aftward loads for all seats. This value will

usually be automatically met by all seats meeting the forward load requirements. Occupant weight shouild be the total weight of the 95th-percentile

crewtiember or trooper as presented in Section 5.8.1.

5.8,2.4 Downward Loads. Human tolerance to vertical impact limits the acceptable forces in the vertical direction for all aircraft seats. The maximum allowable headward acceleration (parallel to the back tangent line) for

seated occupants is on the order of 23 G for durations up to approximately

0.025 sec. Therefore, the 48-G design pulse imposes the requirement for

energy absorption in the vertical direction by some form of load limiting.

The effective weight in the vertical direction of a seat occupant is approximately 8U percent of the occupant's total weight because the lower extremities are partially supported by the floor. The effective occupant weight may

be determined by summing the following:

0 Eighty percent of the occupant's body weight.

a Eighty percent of the weight of the occupant's clothing

(less boots).

0 One hundred percent of the weight of any equipment carried on the

body above knee level. Combat gear is not usually included in the

effective weight of the pilot or copilot (see Section 5.8.1.1).

However, armored seats are often designed for a 95th-percentile male

occupant wearing a chest protector.

The dynamic limit load for the load-limiting system should be established by

use of a load factor (GL) of 14.5. The dynamic limit load is determined by

multiplying the summation of the effective weight of the seat occupant and

the weight of the movable or stroking portion of the seat by 14.5. The resulting dynamic limit load includes the total force resisting the vertical

movement of the seat in a crash; the dynamic limit load of the energyabsorption system, simple friction, and friction due to binding, etc. This

requirement may be difficult to satisfy with a sliding guidance system because the frictional load varies with contact load which, in turn, varies

with the impact load vector direction. Special treatment of sliding surfaces

can reduce this problem. Relatively friction-free rolling and sliding mechanisms have both been used successfully. A rolling mechanism eliminates the

friction prohlem but can introduce a looseness during normal use. This can

be overcome by spring loading the roller joint.

90
The 14.5-G design criterion, taken from Reference 51, considers the dynamic

response of the seat and occupant. The factor of 14.5 was established to

limit the decelerative loading on the seat/occupant system to less than 23 G

for durations in excess of 0.025 sec (the tolerable level for humans as

interpreted from the Eiband data) in crashes that do not exhaust the stroke

of the seat.

Crewseats should be designed to stroke a minimum distance of 12 in. when the

seat is in the lowest position of the adjustment range. This distance is

needed to absorb the residual energy associated with the vertical design

pulse. Further, the load-limiting system should be designed to stroke

through the full distance available,including the vertical adjustment distance. Since a vertical adjustment of ±21⁄2 in. from neutral is typically

required by crewseat specifications, proper design can provide up to 17 in.

of stroke, depending on seat adjustment position. For inclusion of the 5thpercentile female occupant, additional vertical adjustment would be required.

If it is absolutely impossible to obtain a minimum of 12 in. of stroke, a

systems analysis should be used; the godl of the analysis is to show that

occupant protection is equivalent to the system in which the 12-in. stroke is

available.

For retrofit applications, the maximum protection possible should be obtained

in any component being modified, i.e., seats, gear, etc. Separate test crite,,a have been estabi .hed for seats not having the required 12 in. of

stroke and are presented in Section 5.10.2.2.

Since energy-absorbing systems should be designed for dynamic loading, to

obtain the static test loads, dynamic limit loads should be reduced by the

amount due to rate sensitivity of the particular device used. Further, in

the design of the system the desired total resistive load on the seat should

be obtained by summing the resistive load provided by the energy-absorbing

system and the resistive load resulting from friction and/or other mechanisms

unique to the particular system. Thus, the resistive load of the energyabsorbing subsystem must be less than the load required to decelerate the

seat by the amount of the other stroke-resisting variables.

If the energy-absorbing system is to provide only one force setting, the

effective weight of the 50th-percentile occupant from Tables 11 and 12 should

be used for sizing it in order to ensure a tolerable stroke for the majority

of the occupants, not exceeding the stroke limitations of the seat. These

weights are 142.3 and 160.7 lb for pilot/copilot and troop and gunner seats,

respectively.

In order to use the stroke distance available at maximum efficiency, regard- less of occupant weight, a variable-force load-limiting mechanism is desirable. With an infinitely variable force system, the deceleration levels can

be maintained within acceptable limits (if the stroke is not exhausted) for

the full range of occupant weights. Some benefit may also be obtained from a

device that can provide two or more limit loads that can be selected by the

seat occupant. The selection would be made on the basis of seat occupant weight. In operation then, the occupant would be required to select a limit

load by movement of a lever or dial upon entering the seat.

91
It is recommended that at least a dual-level load limiter (preferably three

or more levels) be used to provide maximum protection over the complete

occupant. weight range.

The interaction between the occupant and the movable seat masses increases

with seat mass. Therefore, the movable seat mass should be minimized.

Troop seats should be designed for the maximum stroke feasible to maximize

protection over the large weight range represented by the fully equipped and

lightly equipped occupant. It is recommended that the full 17-in. seat pan

height normally considered desirable from the human engineering standpoint be

used for energy-absorbing stroke. It is further recommended, as a minimum,

that the limit load of the system be sized using the 14.5-G load factor and

the effective weight of the 50th-percentile heavily equipped occupant

(160.7 lb). Variable-level load limiters sized as discussed previously are

also desirable for troop seat.s only if automatically adjusted, since improper

adjustment of such devices can increase the hazard to the occupant.

5.8.2.5 Upward Loads. A capability to withstand a minimum upward load of

8 G is recommended for all aircraft seats. Occupant weight should be that of

the 95th-percentile crewmember or trooper as presented in Section 5.8.1.

5.8.2.6 Lateral Loads and Deformation Reguirements. The lateral load and

deformation requirements for forward- and aft-facing seats are presented in

Figure 35. Two curves are presented. One is for utility and observation

helicopters, and the other is for light fixed-winy aircraft and aitack and

cargo helicopters. The deflections at the seat reference point should be

measured. Occupant weight should be as specified in Section 5.8.1 and should

be that of the 95th-percentile aircrew member or trooper.

Lateral loading in the forward direction (aircraft reference system) on sidefacing seats should be the same as for forward loading (Figure 34), except

that load limiting should be employed.

For crewseats, the lateral deflection should be minimized; however, it is

doubtful if any great stiffness can be achieved in lightweight hardware. It

is believed •d•niiatp, as a design gonal, to attempt to limit the initial deflection to 1 in. with a 2-in. requirement. Because of the possible loading

rate sensitivity of the seat materials, it is felt to be acceptable to allow

analysis of test data to demonstrate compliance. This analysis might include

adjustments of static test data by use of measured or known deflection and

load data from dynamic tests. Further, in cases where wells are provided

under the seats to increase the available stroke distance, the deformation

should be elastic. This will allow the seat to realign itself with the well

prior to entry after reduction of the lateral and longitudinal loads in those

cases where the loads are relieved soon enough.

5.8.3 Other Seats

The requirements presented for crewseats and troop and gunner seats also

apply to passenger seats and any other seat installed in the aircraft for any

purpose. Unique seats installed for special uses zre not to be exempt.

92
SLate'ral

23

Aceptable iailure area

"(Example for forwar d- oD i N lb oz-r af- 1acinq s eat)

"15 ke j c t 101 n- 1 ,~ 15 rC al • ' -Utility and

cc~z 0 observat ion

LLJ helicopters

C_• 1, Li qht fixed-wing, attack, and cargo

I- z

5a (-I, B S s Cul-vt? ...

0

0 0.25 1.0 2.0 3.0 4.0

TOTAL COi{[ROLLFD DEFORMATION (Y), 114.

MEASURED AT SEAT REFERENCE POINT

FIGURE 35. LATERAL SEAT LOAD AND DEFORMATION REQUIREMENTS

FOR ALL TYPES OF ARMY AIRCRAFT.

5.9 PERSONNEL RESTRAINT HARNESS TESTING

The restraint harnesses are to be statically and dynamically tested along

with the seat and/or structure to which they are attached. However, the lap

belt, shoulder straps, and tiedown straps, including all hardware in the load

path, should be statically tested separately to ensure that all components

possess adequate strength and to determine elongation. The strength and

elongation test requirements of restraint system subassmblies are specified

in Table 9.

Specific component tests, including operational tests, are detailed in a

draft military specification (Reference 52). However, all components and

subassemblies should be statically load tested. Each subassembly should be

tested to its full design load to demonstrate its adequacy. Elongation characteristics should be measured to document these data for comparison with requirements and use in systems analyses.

5.10 STRUCTURAL SYSTEM TEST REQUIREMENTS

Both static and dynamic tests are recommended, and it is also recommended

that all seat and litter systems be tested as complete units. This is not to

imply that component tests are not useful; on the contrary, they can be

93
extremely r-seful and should be employed wherever possible to verify required

strengths.

Upon acceptance of prototype systems tested under both static and dynamic conditions, no further tests should be required except for quality assurance.

Major structural design changes in the basic seat system will require static

retesting of the new system to ensure that no loss in strength has been

caused by the design changes. If the changes could affect the energyabsorbing, or stroking, performance of the seat, additional dynamic tests

should also be conducted. Major structural design changes are those changes

involving principal load-carrying members such as floor, bulkhead, or ceiling

tiedown fittings, structural links or assemblies, seat legs, or energyabsorbing systems. Minor changes, such as in ancillary fittings, can be

accepted without a structural test. However, a significant weight increase

such as the addition of personnel or seat armor, would require additional

testing. In summary, changes that increase loading, decrease strength,

produce significant changes in load distribution, or affect the stroking

mechanism will require retesting.

All testing is to be conducted with the seat cushions in place and, for seats

with adjustments, the seats should be in the full-aft position unless another

position is shown to be more critical. The seat vertical position should be

consistent with normal operation (i.e., th. 95th-percentile occupant with the

seat in the full-down adjustment or the 50th-percentile occupant in the neutral position or as most probably used in flight.). All tests should be conducted under simultaneous conditions of floor bucklir and warping or bulkhead

warping, as illustrated in Figure 20. The cumbinatin,, if virpinig conditions

should be that which represents the most critical cas. f'or seat performance,

such as that most likely to impede seat stroking. For .,ample, considering

the combined-load static test (No. 6 in Table 13) of a seat such as thdt shown

in Figure 20, if the lateral load component were applied to the right, the

right-hand track should be warped upward at the forward end (+10 degrees) to

evaluate the possibility of interference with vertical stroke. Also, the seat

should be mounted for testing on actual aircraft hardware, i.e., tracks or

bulkhead fittings.

5.10.1 Static Test Requirements

5.10.1.1 General. Table 13 presents the static test requirements for complete crewseat units per MIL-S-58095. All static tests should be conducted

under simultaneous conditions of floor or bulkhead buckling and warping as

described above. For static testing of trooo/passenger seats the requirements of MIL-S-85510 should be met. The criteria are different, because the

crash environment is usually less severe in the cabin than in the cockpit.

5.10.1.2 Unidirectional Tests. Where separate strength and deformation

requirements have been specified in Table 13 for longitudinal, vertical, and

lateral loading of seats, the loads should be applied separately. Seats must

demonstrate no loss in structural integrity during these tests and should

demonstrate acceptable energy-absorbing capacity.

5.10.1.3 Combined Loads. Seats must demonstrate no significant loss of

structural integrity under conditions of combined loading as shown in

Table 13 and should demonstrate ability to stroke in the vertical direction

with the transverse loads applied.

94
TABLE 13. SEAT DESIGN AND STATIC TEST REQUIREMENTS

Body Weight

Test Loading Direction Minimum Used in Load Seat Weight Deflection

Ref. with Repect to Load FaLtora Determination Used in Load Limitedd

No. Fuselaoe Floor ) lb (kN ) Determination in-_(M)

Undirectional

Loads

1 Forward 35 250 (114) Full 2 (5.1)

2 Aftward 12 250 (114) Full 2 (5.1)

3 Lateralc 20 250 (114) Full 4 (10.2)

4 Downward 25 200 (91) Full No. Reqmt.

(Bottomed)

5 Upward 5 250 (114) Full 2 (5.1)

Combined Loads

6 Combined

Forward 25 250 (114) Full

Lateralc 9 250 (114) Full

Downwardb e 140 (64) Stroking Full

(Stroking) Part Stroke

NOTES:

(a) The aircraft floor or bulkhead shall be deformed prior to the conduct of static tests

and kept deformed throughout load application.

(b) Forward and lateral loads shall be applied prior to downward load applhcation.

(c) The lateral loads shall be appliid in the most critical direction.

(d) Under load at neutral seat reference point..

(e) Static load factor as necessary to meet dynamic test criteria, Figure 37.

(f) Effective weight of a 250 lb (114 kg) occupant.

5.10.1.4 Load Application Method. The test loads should be applied

through a body block (see Section 5.10.1.5) restrained in the seat with the

restraint system. The loads are to be applied at the expected center-ofgravity location of the occupant or occupants of each seat, as illustrated in

Figure 36.

The loads calculated by multiplying the weight of the occupant and equipment

plus the weight of the seat by the required load factor should be applied continuously, or in not more than 2-G increments while the load-deformation performance of the seat is recorded. Maximum loads need not be held for more

than I sec. Th0n maximum load reached, regardless of duration, is to be used

to assess compliance.

On integrally armored crewseats, care should be taken to assure that the

loads are applied proportionally to the proper assembly or test item to

simulate the loads that would typically be carried by the restraint harness

and the seat support structure. In other words, the portion of the load that

95
15.0 IN. S T

(38 CM) STATIC LOAD

S_ APPLICATION

POINT

0 4FORWARD

2.0 +0.1 IN 10.0 _± 0.5 IN. (5. 0. •25 CM)! ' - (25.4 ± 1.3 CM) \

M _j

SI 3.0 IN. (7.6 CM)

RADIUS (TYP) REFERENCE RADIUS

ELLIPSOIDAL CONTOUR 4.0 j 0.5 IN.

(10.2 ± 1.3 CM)

FIGURE 36. STATIC LOAD APPLICATION POINT AND CRITICAL

BODY BLOCK PELVIS GEOMETRY.

could be expected to be restrained by the restraint harness should be applied

to the body block as described above. The portion of the load representing

inertial loading of the movable assembly should be applied separately at the

center of gravity of the appropriate substructure through another provision.

For example, a lever to proportion the load between the body block and movable section of the seat and a sling to apply the appropriate portion of the

load to the bucket can be used. For seats wit) a relatively heavy frame, the

inertial load of the frame can be applied separately at. its appropriate center of gravity. This technique, although adding complexity to the test setup, asst. es that all components in the seat and restraint system assembly

have been tested to their approximate static design loads and that, as far as

a static test simulation can be extended, performance and structural adequacy

have been demonstrated. For lightweight seats (less than approximately 45 lb

for total seat and restraint system), the total load can be applied to the

body block.

As an alternative, static loads may be applied with a centrifuge. In this

case, a dummy, rather than a body block, will be used to simulate occupant

loads.

96
5.10.1.5 Static Load Body Block Or Dummy. Th~e static test loads should

be applied through a body block contoured to approximate a 95th-percentile

occupant seated in a normal flying attitude. The body block should contain

shoulders, neck, and upper legs, and provide for passage of a lap belt tiedown

strap between the legs. The upper legs should be contoured to simulate the

flattened and spread configuration of seated thighs and to allow the proper

location of the buckle. Critical pelvis dimensions are shown in Figure 36.

Buttock contours must be provided to permit proper fit in a contoured seat

pan. The leg stubs should be configured to permit proper seat pan loading as

the body block rotates forward under longitudinal loading; i.e., the leg stubs

should be only long enough to provide a surface to react the downward lap belt

load component, The side view of the buttocks should include an up-curved

surface forward of the ischial tuberosities to allow the forward rotation of

the body block and loading of the shoulder hardness while maintaining the

primary contact between the ischial tuberosities and the seat pan through the

cushions.

5.10.1.6 Deflection Measurements. Deflection should be measured as close

to the seat reference point as possible to prevent seat structure rotational

deformation from influencing the test results. To simplify these measurements, the seat reference point can be projected to the outside of the seat

pan or bucket.

Normally the restraint system will be attached to the seat. However, if a

unique situation should develop in which the only option for increasing crash

resistance is to attach the system (lap belt and shoulder harness) to the

basic aircraft structure rather than to the seat, certain factors should be

considered. First, the forward and lateral deflection requirements of Figures 34 and 35 need not be considered, because the restraint harness limits

torso and seat deflection. Second, the vertical deflection of the seat pan

still must be considered since the downward movement of the seat pan could

cause excessive slack in the restraint harness, or the harness could limit the

stroke of the seat, depending on where the restraint system is ancored.

Neither of these conditions is acceptable in the design.

5.10.1.7 Load Determination. The total load required for all test directions. except. downwar is determined by multiplying the required load factor

from Table 13 by the total of a body weight of 250 lb plus the weight of each

seat. The total load required for the unidrectional downward (bottomed) test

is determined by multiplying the required load factor by the total of an effective body weight of 200 lb plus the weight of each seat. For the combinedload test. the downward (stroking) load required is determined by multiplying the static load factor necessary to meet the dynamic test criteria in Figure 37 by the total of a body weight of 140 lb (average occupant weight less

portion supported by legs rather than seat) plus the weight of the stroking

part of the seat. For centrifuge tests, the dummy weight should be 250 lb for

all tests except the downward tests where it should be 170 lb, and the centripetal acceleration should apply the load factors of Table 13 for at least

i sec.

5.10.1.8 r::Itiple Seats. Multiple-occupancy seats should be fully occupied when tested. If it is determined that the most adverse loading condition

occurs in other than full-occupancy situations, additional tests should be run

for those conditions.

97
G MAX-

-J • -J

0 620

o ai

a 3 I

0 I

SI

I I

ua'

t0 t1 t 2

TIME, SEC

COCKPIT SEATS CABIN SEATS TEST CONFIGURATION PARAMETER

LIMITS LIMITS

S1

DUMMY INERTIAL tI SEC 0.043 0.059

- ODt 2 SEC 0.001 0087

I G MIN 46 32

G MAX 61 37

O" 0V MIN, 50 50

FT/SEC

2

"tI SEC 0.066 C.081

t 2 SEC 0.100 0.127

G MIN 28 22

DUMMY

INERTIAL 0 MAX 33 27 LOAD AV MIN, 50 60 4

FT/SEC

S3&4 _

DUMMY 1I SEC 0.036

INERTIAL

LOAD t 2 SEC 0.061

r MIN 140

G MAX 51

• AV MIN. 42

Fr/SEC

FIGURE 37. DYNAMIC TEST REQUIREMENTS FOR QUALIFICATION.

98
5.10.2 Dynamic Test Requirements

5.10.2.1 Dynamic Test Requirements for Seats Having at Least 12 in. of

Vertical Stroke.

5.10.2.1.1 Crewseats Desi-gned for a Fixed Load. All prototype crewseats

shall meet the requirements of MIL-S-58095. These seats shall be dynamically

tested to the conditions specified in Tests I and 2 of Figure 37. These test

conditions were determined from the design velocity changes presented in Volume II of the Design Guide. Test 1 is required to ensure that the vertical

load-limiting provisions will perform satisfactorily under simultaneous forward and lateral loading conditions. Test 2 is required to ensure that the

seat can resist the loads produced by the design pulse when applied simultaneously in the forward and lateral directions. The actual aircraft seat

attachment hardware shall be used for mounting the seat in the test fixture.

All tests shall be performed with the inertia reel seat in the "autolock"

mode. The seat shall retain the dummy within the confines of the restraint

harness and shall evidence no loss of structural integrity. Any failure of a

restraint system component or of a primary load-carrying structural member of

the seat shall be unacceptable. A primary load-carrying structural member is

defined as a nonredundant member whose failure would allow uncontrolled motion

of the seat and/or potentially injurious impact of the occupant with cockpit

components. Permanent deformations of the structure which do not present a

hazard to the occupant are acceptable. Webbing slippage at adjusters in excess of 1 in. (25.4 mm) is unacceptable. The initial seat height adjustment

shall be set in the mid-position for ,est 1 and in the fullun position for

Test 2. A clothed Hybrid III or VIP-95 95th-percentile dummy weighing 230 lb

(105 kg) shall be used for Tests 1 and 2. For all tests, the dummy's feet

should be secured in a representative anti-torque pedal position.

5.10.2.1.2 Crewseats Designed with an Adjustable Load Attenuation System.

These seats should be dynamically tested to the conditions specified in all

four tests of Figure 37. Test procedure, conditions, and results should be

th. same as rioted above, except as specified in this paragraph. The in;tial

seat height adjustment should be set in the mid-position for all tests except

Test 2, which should be in the full up position. A clothed Hybrid III or

VIP-95 95th-percentile dummy weighing 230 lb (105 kg) should be used for all

tests except Test 3. lest 3 should use a 50th-percentile dummy of Hybrid III

or CFR Title 49, Chapter 5, Part 572, lightly clothed with both arms removed

at the shoulder joints to simulate a 5th-percentile dummy weight. The adjustable attenuation system should be placed in a load setting correslpunding to a

5th-percentile occupant weight for Test 3, and a 95th-percentile occupant

weight for Tests 1, 2, and 4. For Tests 3 and 4, an accelerometer should be

rigidly attached to the lower seat pan centerline surface at a point 5.5 in.

(14 cm) forward of the seat reference point to measure accelerations parallel

to the seat back tangent line. The acceleration measured during lests 3 and

4, should not exceed 23 G for more than 0.025 sec., whnri measured in aLcord

ance with a SAE J211, Class GO instrumentation system, lhis time duration

should be additive, in a cumulative manner, for all acceleration excursions exceeding 23 G. The minimum acceptable seat stroking distance for Tests 3 and 4

should be 9.5 in. (24.1 ram).

5.10.2.1.3 Cabin Seats All prototype truop/passenger seats should mcuc. thie

requirements of MIL-S-855]0 (Reference 53), which requires dynamic testing to

the conditions specified in Tests I arid 2 uf ligure 37, using a clad 50th

99
percentile dummy (Reference 36) it, Test 1 and a clad 95th-percentile dummy in

Test 2. Dynamic testing of multipla-occupant seats shnuld be per'ormed with

the maximum number of occupan's specified for the test seat. Additional tests

should be run if it is determined that the most adverse loading condition

occurs in other than full-occupancy situations or that occupant size is a factor. For both tests of Figure 37, adjustable seats should be adjusted to the

full-aft and kip position of the adjustment range. Plastic deformation of the

seat is permissible; however, structural integrity must be maintained in all

tests. For Test 1, the seat should limit the acceleration as measured in the

pelvis of the dummy to values which ensure that the 50th-percentile clothed

seat-system occupant (see Section 5.8.1) will not experience vertical, +Gz,

accelerations in excess of human tolerance as defined in Sections 5.3 and 5.9

of Volume II (see Figure 9 herein). The roll direction (10 degrees right or

left) for Test 1 should be selected to produce the more critical loading for

the specific seat design.

When determining compliance of the achieved test pulse with the dynamic test

requirements of Figure 37:

1. Determine the maximum acceleration and construct the onset slope for

the test pulse by the method explained in Section 5.10.3.

2. Compare the achieved onset and peak acceleration of the test pulse

with those allowed and presented in Figure 37. lhe achieved onset

slope should lie between the minimum and maximum onset slopes using

the values of ti and t 2 listed in Figure 37 for the specific test

conditions. Hi, ,axiriiuii acceleration should also fall between the

upper and lower limits allowed.

3. IntegraLe the actual acceleration-time curve of the test pulse and

establish the achieved velocity change. The velocity change achieved

should be equal to or greater than that tabulated for the specific

test conditions.

5.10.2.2 SpecialQynamaic lest Re~ulrements for Seats Having Less Than 12 in.

of Vertical Stroke. In the event that the application of a systems approach

permits the seat to have less than 12-in,. minimum vertical stroke or retrofit

restraints preclude available room, additional requirements are made of the

dynariiic Lesting. first, it would be desirable to periorm a full-scale crash

test with the test specimen, including all assemblies involved in the energyabsorbing process. 1his would include a section of the fuselage, landing

gear, and the seat or seats. This approach is totally acceptable for demonstrating the dynamic response anrd acceptability of the system.

Since cost associated with the type of system testing described above is

usually prohibitive, a different approac(h is acceptable. This approach in

clides dynamically testing the seat only, as is done for systems with at least

12 in. of stroke, but modifying the input pulse to represent the energyab-or)irbig proLesses of the gear and fuselagc. An example of such a modified

te..t pulse is presented in l iguru 38. lhe initial plateau (t1 to t')

represents the acceleration-time history created by stroking of the landing

gear. 1hu sharp increase in acceleration at t' relates to fuselage impact,

and the pulsue beyond t ) represents thl, crushhirni of the stiffer fuselage

100
G2

z I

0 I

<

cc

-J

&U

G FUELG

G 1 - Jl :USE AGE m - S LANDING GEAR STROKING CRUSHING

to t1i t6 t1

TIME (SEC)

FIGURE 38. EXAMPLE OF INPUT PULSE FOR SEATS HAVING

LESS THAN 12 IN. OF STROKE.

section. The velocity change under the pulse should be the same as identified

for the particular crash force direction for other established tests

(50 ft/sec for Test No. I or No. 2 of Figure 37).

It will be difficult to determine accurate dynamic crush characteristics of

the various portions of the system to enable establ;shment of a representative, and thus acceptable, test pulse. The best analytical techniques,

supported by test data, should be used for determining the properties of the

fuselage. Since drop tests of landing year are required, a much more accurate

approach exists for obtaining the landing gear influence on the pulse. Seat

testing should await completion of landing gear tests so that the results can

be used to establish the initial plateau (or other shape) between tI and t'

of the input pulse.

Typically the landing gear will stroke at loads below those required to stroke

the seat; therefore, much of the kinetic energy of the occupant and 5eat will

be absorbed prior to fuselage impact. If the systems analysis is accurate,

the energy-absorbing capacity of the seat will be sufficient to absorb the residual energy at limit loads tolerable to the occupant.

101
Since each system may display different characteristics, it is not appropriate

to present in this document specific quantitative limits for use in evaluating

the acceptability of the test pulse. However, the same general approach and

tolerances already presented for the standard pulse apply and should be used.

The technique described in Section 5.10.2.1 for establishing compliance with

the required test pulse applies directly to the portion of the special test

pulse following to.

5.10.3 Data Acaisition and Reduction

Data acquisition and reduction should comply with the requirements of SAE J211

(Reference 54) for measurements on anthropomorphic dummies and structures.

Dynamic test data must usually be smoothed by filtering out high-frequency

data and/or noise to be useful. This is especially true if it is to be

sampled and digitized. It is good practice to use filtering procedures common

to other test laboratories, as this eases valid comparison of results. The

suggested criteria for data filtering are found in Figure 1 and Table 1 of

SAL J211. These are reproduced in Table 14 for convenience. Data should be

visually examined in the unfiltered state to assure that saturation or other

distortion did not occur.

TABLE 14. DATA CHANNELS

Re,.ponse Rangei2)

lest Measurement Channel Class Hz

Dummy

Head acceleration 1000 O.i - 1000

Chest ar.celeratlon 180 0.] - 180

Femur force 600 0.1 - bOO

Restraint systenm loads bO 0.1 - 60

Sled ot vehncle dcceiera.iuri 6u 0.1 - .0

(1) Except for coompunent analysis u.•u Charinni Class 600 and for

irnteyratiun for velcity use Channel Class 180.

I2) lit response 4_ 1W dl3 at low end to 1 1/2 -1 dB at high end.

Iilter rolluff chaiautvrnstics above high end are defined in

SAL J211.

Ir,struments for dynamic measurements must have the proper frequency response

rang(. to prevent. distortion of the data. In addition to adequate highfrequency response, response to 0 11 is needed to prevent distortion of

low-frequency data which is also typ[ically found in crash data. Therefore,

piuzoresistive or wire strain gage devices are preferred over piezoelectric

devices. Instrumreits should also be calibrated over the frequency range of

interest. A centrifuge calibration of an accelerometer, for exampie, really

cal ibrates the device under static G loading conditions. That calibration

102
may not accurately represent the performance of the device under dynamic conditions. A dynamic calibration over the entire frequency range of interest

is preferred.

Data should be presented in both analog and tabular form in compliance with

the sign convention shown in Figure 5. Impact velocity should be determined

and recorded for the test platform or vehicle. In the analysis of the data,

velocity change should be computed through either electronic means or graphically Oith a planimeter by integrating the area under the measured acceler3tion-time trace.

The method recommended for use in establishing the acceptability of the pulse (see Section 5.10.2) and to determine other parameters associated with the

data is similar to that presented in MIL-S-9479(USAF) (Reference 55). Parameters such as rise time, onset slope, and acceleration plateau duration may be obtained using the following graphic approximation technique shown in

Figure 39:

a Locate the calibration baseline.

* Determine the maximum (Gp) acceleration magnitude.

* Construct a reference line parallel to the calibration baseline at a

magnitude equal to 10 percent of the peak acceleration (G ). The

first and last intersections of this line with the accele~ation-time

plot defines points I and 2.

* Construct a second reference line parallel to the calibration baseline at a magnitude equal to 90 percent of the peak acceleration.

The first and last intersections of this line with the accelerationtime plot define points 3 and 4.

* Some practical judgment may be required for selection of the first

and last intersections depending on the degree of noise apparent in

the data. Significant tendencies are important, not noise.

* Construct the onset line defined by a straight line Lhruugh puiniLs I

and 3.

* If desired, construct the offset line defined by a straight line

through points 2 and 4.

* If desired, construct a line parallel to the calibration baseline,

through the peak acceleration. The time interval defined by the intersections of this line with the constructed onset and offset lines

(points 5 and 6) is the plateau duration (vt).

* Locate the intersection of the constructed onset line with the calibration baseline (point 7). The time interval defined by points 7

and 5 is the rise time (tr -- to). Referring to Figure 37, the

rise time should be greater than t1 but less than tý when determining compliance with dynamic test requirements. Point 7 in Figure 39 is the initial time to in Figure 37.

103
0.9 Gp 3 4

Calibration

baseline 0.___ 2

FIGURE 39. GRAPHIC APPROXIMATION EXAMPLE. (REFERENCE 55)

5.10.4 Seat Component Attachment

Since components that break free during a crash can become lethal missiles,

it is recommended that attachment strengths be consistent with those specified for ancillary equipment mounted to the seat (see Volume III). Therefore, static attachment strengths for components, e.g., armored panels,

should be as follows:

Downward: 50 G Aftward: 15 G

Upward: 10 G Lateral: 25 G

Forward: 35 G

5.11 RETROFIT FOR SEATING SYSTEMS

5.11.1 General

If a retrofit effort is to install crash-resistant seats in an existing airFrame, complex interface problems may result. This is because the seat attachment points on the airframe were not designed for the loads which will be

imposed by a crash-resistant seat. The first, and preferred, approach is to

calculate the loads required to support a crash-resistant seat and then determine how the floor or bulkhead should be modified to support those loads.

Seat design will then proceed as discussed in previous sections. If the

impact velocities of the retrofitted aircraft are significantly different

from the recommended standards presented herein, then more representative

velocities should be used to design the retrofitted seats.

104
5.11.2 Forward-Load-Limiting Seats

If, for any reason, the aircraft attachments cannot be modified to support

the loads applied by a crash-resistant seat, then another approach is possible. That is to design features into the seat which will permit it to limit

loads applied to the aircraft. It can be accomplished through controlled deformation of the seat structure. The technique has been used for crewseats

for both the SH-3 and CH-53 helicopters. For each of these aircraft, crewseats were designed which limited loads in the forward and lateral directions

as well as the downward direction. The forward and lateral load limiting protects the attachment structure and has nothing to do with human tolerance.

The downward load limiting is determined by human tolerance considerations,

as discussed in previous sections.

Figure 40 shows a sketch of the CH-53 crew seat. The rear struts are energyabsorbing devices which will elongate at a fixed constant load. This permits the center of gravity of the seat occupant system to move forward relative to

the floor attachment and limits the attachment forces. The back view of the

seat in Figure 40 shows the high elongation diagonal braces which allow the

seat and occupant cg to move sideways at a controlled load. Thpse braces simply employ the plastic stretching of metal. The seat designed for the SH-3

uses the same techniques. These seat systems are further described in References 19 and 20.

S1

ARMORED WING PANEL AR..ORE. BUCKET

VERTICL /WING ~ PANEL

ENERGY ABSORBERS 'h\\' GIDE T'!BES 'r

LINEAR BEARING *~ ~jt

FIGULRE AR ST. H-BEARU ENER T ABSORBERS 0

DIAGON.AL BO.. -.

FIGURE 40. CH-53 CREW SEAT.

105
5.12 LITTER STRENGTH AND DEFORMATION REQUIREMENTS

5.12.1 General1

The ultimate vertical strength of existing litters with a 200-lb occupant and

a total system weight of 250 lb (see Section 5.12.2) is about 13 G. Since

the desired decelerative loads to be imposed on these litters exceed 13 G,

special techniques must be used to limit the deflection and to support some

of the occupant load.

Lateral orientation im1 the aircraft is preferred because of the characteristics of existing restraint systems used on litters which provide more support when loaded laterally than when loaded longitudinally.

5.12.2 Recommended Occupant Weights for Litter Design

The litter strength and deformation requirements defined below are based on a

200-lb, 95th-percentile litter occupant with 20 lb of clothing and personal

gear, a 10-lb splint or cast, and 20 lb of litter and support bracket weight

for a total weight of 250 lb (the weight of a litter and patient as specified

in MIL-A-8865 (ASG), Reference 56).

5.12.3 Vertical Loads

5.12.3.1 Downward Loads. In the case of litter systems, human tolerance

is not the limiting case in th. vertical direction. The loads would be applied in a transverse direction to the body of a litter occupant. However,

design to the 45-G human tolerance level is impractical due to the strength

requirements for litters and for the basic structure to support the litter

systems.

Litters are either hung from the ceiling or suppcrted at the floor. In

either case, the input deceleration pulses are the same as for floor- or

bulkhead-mounted seats. Litters should not be suspended from the overhead

structure unless it is capable of sustaining, with minimum deformation, the

downward loads from the tiers of litters. Therefore, in the design of an

efficient system, intentional load limiting should be related to the floor

pulse.

The vertical strength and deformation requirements for a litter system are detailed in Figure 41. This curve is read in the identical manner as the seat

load-deflection curve shown in Figure 34. The load factors in units of G are

based on the summation of the weights of the occupant plus clothing, personal

gear, splint or cast, and the weight of the litter and attachment brackets

for a total of 250 lb as described in Section 5.12.2. The curve of Figure 41

is based on the assumption that 3 or 4 in. of vertical deflection will occur

at the midpoint of the litter. In the unlikely event that a rigid litter is

used, an additional 2 in. of deflection should be added to the curve. The

deflection curve is limited to 6 in., because a large deflection occurring on

one corner of the litter due to an asymmetric loading could cause ejection of

106
the litter occupant. A larger energy-absorbing stroke can be used effectively if a mechanism is included in the system to control the amount of tilt

allowed. For example, a system mechanism could be designed that forced all

four corners of the litter to stroke the same distance (within elastic

limits) thus achieving this goal.

Aircraft ceiling

Downward

-r--777- 7 ,7.7 7/7 7777z 7

Aircraft floor

SACCEPTABLE

AAll•M ýC Pi• : [REJECT ION FAILURE 0 2- ADEK7A AREA

-4-

BASE AREA

10 " - -

0

*H5

4_1

0 1 2 3 4 5 6

Total controlled

deformation (z), in.

*G value based on 250-lb per litter

position.

FIGURE 41. LITTER DOWNWARD LOAD AND DEFLECTION REQUIREMENTS.

The additional problem associated with inadequate litter strength must be

dealt with in the design of litter systems. The curve of Figure 41 assumes a

litter capable of at least 17 G with a maximum of 25 G. If the existing

107
litter is used, then a pan, net, or other device should be included under the

litter to catch and support the litter occupant if the litter fails.

Actually, the device should limit the deflection to a value less than that

required to fail the litter and should stroke with the litter. If all of

these provisions are included, i.e., a rigid new litter or old litter with

supporting pan underneath, together with the tilt-limiting mechanims, then

the stroke can be extended to 12 in. at a 17-G limit-load factor. The loaddeformation curve of Figure 41 would be extended at 17 G to 12 in. of stroke.

5.12.3.2 Upward Loads. All litter systems should be capable of withstanding a minimum upward load of 8 G.

5.12.4 Lateral and Longitudinal Loads

Litter systems for all military aircraft should be designed to withstand the

load and deformation requirements indicated in Figure 42 in all radials of

the lateral/longitudinal plane. The litter lateral loads are made equal to

the longitudinal loads because the litters may be oriented in either direction depending upon the aircraft.

3030 Acceptable failure area

IiA cceptableJ

- 2 5 . ' "__-

S...nac ceptable

20 .- /__Rejection ara_

io _ _ _ __ Base 1 _ _ _ _ _ _ _ curve

0

4 4

0

0 1 2 3 4 5 6

Total controlled forward*" deformation of litter bed, in.

*G value based on 250-lb per litter position.

**Forward is the direction towards the nose of the aircraft

regardless of litter orientation in the aircraft.

FIGURE 42. LITTER FORWARD OR LATERAL LOAD AND DEFLECTION

REQUIREMENTS FOR ALL TYPES OF ARMY AIRCRAFT.

108
The 20-G acceptable load level indicated in Figure 42 is predicated on the

tolerance to acceleration of an individual restrained by straps on existing

"table top" litters. If litters and allied restraint harnesses are designed

for improved crash resistance, the 20-G load should be increased to 25 G.

Acceptable or nonacceptable load-deformation characteristics are read from

Figure 42 in the identical manner as the readings from Figures 34 and 35 for

seats. The deformation is measured with respect to the aircraft floor along

the longitudinal axis toward the nose of the aircraft, regardless of litter

orientation.

5.12.5 Litter Restraint Harness Testing

The restraint used in existing military litters consists of two straps

wrapped around the litter. These straps should withstand a straight tensile

minimum load of 2000 lb (4000-lb loop strength). The maximum elongation

should not be more than 3.0 in. under the straight pull (end-to-end) test on

a minimum strap length of 48 in. Elongation is restricted for litter belts

in order to minimize dynamic overshoot.

5.12.6 Litter System Test Requirements

5.12.6.1 S!,vtic Test Requirements

5.12.6.1.1 General. Table 15 presents the static test requirements for

complete litter systems. Since previous studies have shown that existing

litters will not withstand the loads as specified in this chapter, the

assumption must be made that a litter of sufficient strength will be

developed prior to implementing these recommendations. If a pan or net to

catch the litter occupant is included in the system, it should also be

included in the static testing to demonstrate its adequacy.

TABLE 15. LITTER SYSTEM STATIC TEST REQUIREMENTS

Test Loading Direction

Ref- With Respect to Deformation

No. Fuselage Floor Load Required Requirements

1 Forward See Figure 42 See Figure 42

2 Lateral See Figure 42 See Figure 42

3 Downward See Figure 41 See Figure 41

4 Upward 8 G No requirement

5 Combined loading

Downward plus See Figure 41 See Figure 41

transverse load

along any radial

in the x, y plane

of the aircraft See Figure 42 See Figure 42

109
5.12.6.1.2 Unidirectional Tests. The test loads for forward, lateral,

and downward loading of litter systems as presented in Table 15 should be

applied separately.

5.12.6.1.3 Combined Loads. Litter systems must demonstrate no loss of system integrity under conditions of combined loads as specified in Table 15.

5.12.6.1.4 Point of Load Application. The loads should be applied through

a body block that simulates a supine occupant.

5.12.6.1.4.1 Forward (Longitudinal) - Lateral Tests. For systems using

the existing litter, a rigid simulated litter maý be substituted for the actual littev. This will enable application of equal loads at all attachment

points between the litter and the suspension system and allow testing of the

suspension system. The rigid litter substitution does not apply if the

litter system has adequate strength to take the loads.

5.12.6.1.4.2 Downward and Upward Tests. Downward and upward loads may be

applied to each vertical suspension point separately. If the suspension

system has the tilt-limiting features, and the litter strength is adequate,

then the load should be applied at the center of gravity of the body block.

5.12.6.1,5 Deflection MHdsurements. Downward, forward (longitudinal), and

lateral deflections should be measured at the bracket attaching the litter to

the suspension system.

5.12.6.1.6 Load Determination. The test load should be determined by

multiplying the required load factor (G) as specified in Table 15 by 250 lb.

5.12.6.2 Lit System Dynamic Test Requirements. A single test to evaluate the ver, load-limiting system is required. Litter systems with

95th-percenti_- anthropomorphic dummies and 30 lb of additional w,'ight

(250-lb total) in each litter should be subjected to a triangular acceleration pulse of 48-G peak and 0.054-sec duration (42-ft/sec velocilj change).

The same test pulse tolerances, data, handling, and processing requirements

as presented for the seats in Section 5.10.2 apply. At least three accelerometers should be placed in the dummy; one in the head, one in the chest, and

one in the pelvic region. The instruments shoild be positioned to sense accelerations in the vertical directions (x-axis of the supine occupant, zdirection relative to the aircraft). The input acceleration-time pulse also

should be measured. It is advisable to use redundant accelerometers to sense

the input pulse to assure acquisition of the needed impact environment data.

5.13 DELETHALIZATION OF COCKPIT AND CABIN INTERIORS

.J.13.1 General

ihe kinematics of body action associated with aircraft crash impacts are

quite violent, even in accidents of moderate severity. The occupant's

immediate environment should be designed so that, when the body parts do

flail and contact rigid or semi-rigid structures, injury potential is

minimized.

110
Several approaches are available to alleviate potential secondary impact

problems. The most direct approach, which should be taken if practical, is

to relocate the hazardous structure or object out of the occupant's reach.

Such action is normally subject to trade-offs between safety and operational

or human engineering considerations. If relocation is not a viable alternative, the hazard might be reduced by mounting the offending structure on

frangible or energy-absorbing supports and applying a padding material to

distribute the contact force over a larger area on the body member.

5.13.2 Occupant Strike Envelopes

5.13.2.1 Full Restraint. Body extremity strike envelopes are presented in

Figures 43 through 45 for a 95th-percentile Army aviator wearing a restraint

system that meets the requirements of MIL-S-58095 (Reference 15). The restraint system consists of a lap belt, lap belt tiedown strap, and two

shoulder straps. The forward motion shown in Figures 43 and 44 was obtained

from a test utilizing a 95th-percentile anthropomorphic dummy subjected to a

spineward (-Gx) acceleration of 30 G. The lateral motion is based on an

extrapolation of data from the same 30-G test. In positions where an occupant is expected to wear a helmet, the helmet dimensions must be added to the

envelope of head motion.

__---___

10 In.

HELREST .'NE FOR COCKPIT

AIRCRAFT FLOOR LINE FOR TROOP COMPARTUENT

FIGURE 43. FULL-RESTRAINT EXTREMITY STRIKE

ENVELOPE - SIDE VIEW.

11i
_TT

i.n.," 10o ..

_ _ _.-. _ _ ,_ 

I " ... .... _ J ... _ Q_ es-t- . . ..

FIGURE 44. FULL-RESTRAINT EXTREMITY STRIKE

ENVELOPE -- TOP VIEW.

/

.- - I - ._,1_- - ('.-----'•-f ___o;;i

ENVELOPE-- FRON VIEW.

. .... S..1 1 2
5.13.2.2 LaD-Belt-Only Restraint. Although upper torso restraint is required in new Army aircraft, strike envelopes for a 95th-percentile aviator

wearing a lap-belt-only restraint are presented in Figures 46 through 48 for

possible use. They are based on 4-G accelerations and 4 in. of torso movemeot away from the seat laterally and forward. In positions where an occupant is expected to wear a helmet, the helmet dimensions must be added to the

envelope of head motion.

to....

, 11

. ... ra t floio lin, fol trv. L.OflJ'4um)•r tIncit.Il

FIGURE 46t LAP-RELT-ONLY EXTREMITY STRIKE

ENVELOPE - SIDE VIEW.

5.13.2.3 Seat Orientation. The strike envelopes of Figures 43 through

48 apply to all seat orientations.

5.13.3 Head Strike Envelope in Stroking Seats

The head strike envelope for a stroking energy-absorbing seat is obviously exaggerated relative to the above diagrams. Reference 57 describes some simulations which were performed to evaluate the head strike envelope in this situation. Additional information may be found in Volume IV.

5.13.4 Environmental Hazards

5.13.4,1 Primary Hazards. The primary environmental hazards are those

rigid or semirigid structural members within the extremity envelope of the

head and chest. Since the upper torso, and particularly the head, is the

most vulnerable part of the body, maximum protection must be provided within

its strike envelope.

113
FIGURE 47. LAP-BELT-ONLY EXTREMNITY STRIKE

ENVELOPE TOP VIEbi.

-44

FIGURE 48. LAP-BELT-ONLY EXTREMITY STRIKE

ENVELOPE - FRONT VIEW.

114
5.13.4.2 SecondaryHazards. Secondary environmental hazards are those

that could result in trapping or injuring the lower extremities to the extent

that one's ability to rapidly escape would be compromised. Areas within the

lower extremity strike envelope must also include ample protective design.

5.13.4.3 Tertiary Hazards. Tertiary environmental hazards are those rigid

and semirigid structural members that could cause injury to flailing upper

limbs to an extent that could reduce an occupant's ability to operate escape

hatches or perform other essential tasks.

5.13.5 Head Impact Hazards

5.13.5.1 Geometr1y of Probable Head Impact Surfaces. Typical contact hazards in the cockpit area include window and door frames, consoles, controls

and control columns, seat backs, electrical junction boxes, glare shields,

and instrument panels. Contact hazards commonly found in aircraft cabin

areas include window and door frames, seats, and fuselage stru(ture. Use of

suitable energy absorbing padding materials, frangible breakaway panels,

smooth contoured surfaces, or ductile materials in the typical hazard areas

mentioned is recommended to reduce the injury potential of occupied areas.

5.13.5.2 Tolerance to Head Impiact. Protection of the head in the form of

protective helmets and energy-absorbing structure and padding in the occu

pant's immediate environment is essential.

lolerance levels fur head impact are discussed in detail in Volume If, and

the reader should refer there fur an understanding of the problem. However,

for the case of forehead impact on a flat surface, which is pertinent to the

discussion of this section, the most widely accepted collection of tolerance

data is represented in the tolerance curve of figure 49 (Reference 58).

5.13.5.3 'rest Procedures. The simplest test procedure for evaluating

the effectiveness of protective structure and padding in prevenLing seriou-s

head injury makes use of an instrumented headform. The headform, equipped

with an accelerometer, can be propelled by a ram, dropped, er swung on a

pendulum to imrnacut the surface to be evaluated. lhis procedure is described

in SAL J921 (Referenice 59). The n, easured acceleration pulse can be averaged

for comparison with the Wayne State Iolerance Curve, or intugrated to compute

a Severity Index, as discussed in Section 5.4.1 of VolumiL I.

Another approach is to use simulations or tests of the entire occupant seat

system. Section 11.4.5 of Volunime IV presents information on this approach.

5.13.6 Isltrument Panel Structure Prpximnty

In most aircraft coc.kpits, the i nstrurnent panel and its supporting structure

are placed directly above the pilot's lower leg',, lhe danrge;r of impact iromn

this proximity dictates that designers consider using suitable energyabsorbing padding materials, frangible breakaway panels, or ductile panel

materials fur structure within the lower ley strike envelope.

As discussed in Section 11.5.1 of Volune IV, the use uf a fiberylass instrumerit glare shiield may be use'd as arn alternative to padding or to provide

additional protection from protruding instrument!s.

115
600 -

0

400

C: DANGEROUS TO LIFE

W

-IJ

uJ 300 - .---- .

'U

> 200

TO LIFE

0 V - 1- 4--J

0 2 4 6 8 10 1230100

TIME DURATION OF EFFECTIVE

ACCELERATiON (MSEC)

FIGURE 49. WAYNE STATE TOLERANCE CURVE FOR THE HUMAN

BRAIN IN FOREHEAD IMIPACTS AGAINST PLANE,

UNYIELDING SURFACES. (REFERENCE 58)

5.13.7 Rudder Pedal Configuration

Rudder pedals should be capable of supporting both the ball of the foot and

the heel, and provide a surrounding structure of sufficient strength to

prevent crushing and trapping of the lower 1 imbs. Ihe geometry required by

MIL-SlD-1290 (Reference 1) to prevent entrapment of feet is illustrated in

Iigure 50.

5.13.8 (ontrols and Control Columns

Control colurrins located in front of flight crew stations can present a seriouS hazard to crewme"bers i, they fail at any appreciable distance above the

aircraft. flour. 1he failure should occur in the form of a clean break,

leavin(q no jagged or torn udges. Control columns that pass longitudinally

through thv i nstruwmcnrt pdriel are riot. recommnded sirice these tend to impal e

thi. crewmvimbers in severe lonjitudinal impacts. However, where they are used

they should bu equipped with a frangible or energy absorbing section simnilar

to automotive steering Lolulni,.

]16
41 A 13 - --- C *

/i

/¥

NEW:

Dirnen~ion3 A, B, and C DIUut be CitILU-1

less than 2 in. or mr-e than 6 in.

FIGURE 50. ANTITORQUE, OR RUDDER, PEDAL GEOMETRY

TO PREVENT ENTRAPMENT OF FEET.

The cyclic control stick is an examplc of a lethal object which may be involved in head impacts. This hazard may be increased if stroking energy- absorbing .Ueats are installed, because, as the seat strokes, the cLewittber's

head comes closer to the stick. Section ]1.7 of Volume IV discusses means of delethalizing the stick.

5.13.9 Sighti GUREnd Vi0. ionic. systemsD

Delethalization of the copilot/yunne2r (CPG) station of an attack or scout hel- icopter equipped with a weapon epghting optical relay tube (01c) man present

a divvficut design problem. The cockpit should be designed to minimize n he

probability of the CPG head/neck striking the, ORl and minimize inju~ry if the CaG should strike hetal ORT, for both the "head-up" and "head-down" CPG positions. Some of the optiont available to the designer V iven this task are:

a ORT designce pRelo.cation - Consideration should be given to reducing

occupant strike hazards by moving the OR] farther aw&y from the

CPG.

117
* Restraint System - The rqstraint system of Figure 24 would offer improved upper torso restraint, particularly when combined with the

power-haulback inertia reel.

* Inflatable Restraint - Consideration should be given to the inflatable restraint system (IBAHRS) discussed in Section 5.7.2.4. This

type of restraint harness can prevent injury to the CPG in both the

erect and head-down position by reducing slack, supporting the head,

and increasing the surface area of the body over which the harness

reacts.

* Franqible/Breakaway Features - ORT or ORI components designed to be

frangible should break away at a total force not to exceed 300 lb.

For the frangible ORT, this force should be applied along any direction of loading within the plane normal to the axis of the ORT, as

well as along the axis of the ORT. Breakaway point(s) of the ORT

should be outside the head strike envelope.

* Collapsible Features - If the ORT is designed to collapse in order to

avoid injuring the CPG, the collapse load along the axis of the ORT

should not exceed 300 lb. Figure 51 illustrates one crushable sight

eyepiece concept (from Reference 60). Two advantages of the crushable sight eyepiece are that it is always available and, it should

function regardless of head location. A helmet crash-absorber pad would attenuate crash loads to the helmet when available crushing is

exlJýlueu.

* Power-Haulback Inertia Reel .. IBI__R). - On the basis of Air Force testing accomplished for the development of PHBIR, the retraction time is

0.3 to 0.4 sec, which is too slow for effectiveness in most crashes.

If this time were reduced, the retraction velocity of the torso would

have to be increased considerably over the current limit of 9 ft/sec.

A retraction velocity greater than this is not recommended due to the

lack of human tolerance data on this type of loading. In a crash

with a single pulse of 30-G prak and 50-ft/sec velocity change, the

retraction velocity should be approximately 25 ft/sec; therefore, the

known tolerance limits would be exceeded at the higher velocity. In

summary, the PHBIR, as currently qualified under both Air Force and

Navy military specifications, requires excessive time to position the

torso by crash sensing. To be fully effective, the system should

move the torso into position in approximately 0.06 sec, but the resulting acceleration would exceed known human tolerance limits. The

primary crash- resistance advantage of the PHBIR would be as a manually activated tightening device for the head-up CPG position; the

PHUIR offers only limited advantage for the headdown CPG position.

5.13.10 Enerqy-Absorbin Requirements for Cockpit and Cabin Interiors

5.13.10.1 General. 'o minimize occupant injury, the acceleration experienced during secondary impacts of the occupant with surrounding structures

must be reduced to a tolerable level. The areas of contact to be considered

for ernergy absorption include instrument panels, glare shields, other interior surfaces within the occupant's strike envelope, and suat cushions. A

padding material should not only reduc, , the decelerative force exerted on anr

118
Normal operation

---- SPH-4 helmet , reference Helmet crash- -- / Henmetcrgy abso Pressure support energy absorber / -ae prxmtl

area approximately

Crash energy- 6 in.

absorbing material /7

PreladedInner contours slip rings - ,•

OoCrush stroke o face mnask

optical relay Outer contour

tube _ tube < -face mask

20-MM

eye relief

After crash stroke

Helmet crashenergy absorber

Yiele

material " /

Deformed tube

FIGURE 51. CRUSHABLE EYEPIECE CONCEPT. (REFERENCE 60)

impacting body segment but should distribute the load in order to produce a

more uniform pressure of safe magnitude.

In order to prevent head injury, materials must be carefully selected to absorb and attenuate the energy of impact. The material must reduce the level

of acceleration, the rate of onset, and the amount of energy transmitted to

the head.

5.13.10.2 Padding Material Properties. The selection of a foam material

for vehicle energy-absorbing applications involves an evaluation of its processability; its mechanical, thermal, and chemical properties; as well as its

119
cost. Along with the primary foam materials, the characteristics of adhesives and surface coatings must be considered, particularly with respect to

emission of smoke and toxic vapors. The characteristics of suitable materials for such use are listed below:

0 Adaptability and ease of 0 Nontoxic fume generation

processing . Favorable flammability

* High energy dissipation rating

* Effective load distribution • Minimal smoke generation

* Low rebound e Durability and long life

• Temperature insensitivity • Cost competitive

* Resistance to chemicals, oil, * Aesthetically acceptable.

ultraviolet radiation, and

sunlight

5.13.10.3 Standard Test Methods. ASTM standard test procedures are widely used by manufacturers to specify various properties of a particular type

of material. Table 16 summarizes ASTM test methods and specifications for

flexible cellular plastics that provide a basis for comparison of materials.

Here it may be noted that most ASTM tests involve simple tests, whereas the

operati onal tnvitronment involtvest dynami-c. loading, and more comp!ex conditions.

In particular, ASTM D 1564-71 describes "Standard Methods of Testing Flexible

Cellular Materials-Slab Urethane Foam" (Reference 61). Among other tests,

there are compression-set and load-deflection tests.

The above tests provide results that specify the material, but do not necessarily portray its performance under actual impact situations. A simple dynamic drop test, such as ASTM D1596-64 (1976), "Standard Test Method for

Shock-Absorbing Characteristics of Package Cushioning Materials" (Reference 62), more closely simulates actual impact conditions.

Other standard test procedures include SAE J815, "Load Deflection Testing of

Urethane Foams for Automotive Seating" (Reference 63), which points out the

factors of interest in testing materials for vehicle seat c'shions: the

thickness of the padding under the average passenger load, a measurement that

indicates the initial softness, and a measurement that indicates resiliency.

Also, SAE J388, "Dynamic Flex Fatigue Test for Slab Urethane Foam" (Reference 64), describes procedures for evaluating the loss of thickness and the

amount of structural breakdown of slab urethane foam seating materials.

SAE J921, "Motor Vehicle Instrument Panel Laboratory Impact Test Procedure -

Head Area," describes a test procedure for evaluating the heal impact characteristics of such areas as instrument panels (Reference 59).

5.13.10.4 Acceptable Stress-Strain Characteristics. Energy-absorbing materials with stress-strain curves that fall between the limits shown in

Figure 52 will offer reasonable survival potential for head impacts at

velocities of up to 22 ft/sec where a padding thickness of 2.0 in. is used.

120
TABLE 16. SUMMARY OF ASTM TEST METHODS AND SPECIFICATIONS FOR FLEXIBLE

CELLULAR PLASTICS (REFERENCE 61)

D1564-71* Testing Flexible Cellular Materials - Slab Urethane Foam

D1667-76* Specificatiorn for Flexible Cellular Materials - Vinyl Chloride

Polymers and Copolymers (Closed-Cell Sponge)

D1565-76* Specifications for Flexible Cellular Materials - Vinyl Chloride

Polymers and Copolymers (Open-Cell Foam)

D1055-69" Specification for Flexible Cellular Materials - Latex Foam

(1975)

01056-73" Specification for Flexible Cellular Materials - Sponge or

Expanded Rubber

D3575-77 Testing Flexible Cellular Materials Made from Olefin Plastics

D1596-64* Test for Shock-Absorbing Characteristics of Package Cushioning

(1976) Materials

02221-66' Test for Creep Properties of Package Cushioning Materials

(1973)

D1372-64' Testing Package Cushioning Materials

(1976)

0696-70* Test for Coefficient of Linear Thermal Expansion of Plastics

E143-61" Test for Shear Modulus at Room Temperature

(1972)

0412-75" Tests for Rubber Properties in Tension

01433-76" Test for Rate of Burning and/or Extent and Time of Burninq of

Flexible Thin Plastic Sheeting Supported on a 45-degree Incline

01692-76 Test for Rate of Burning and/or Extent and Time of Burning of

Cell ar Plastics Using a Specimen Supported by a Horizontal

Screen

*Indicates that the standard has been approved as American National Standard

by the American National Standards Institute,

121
20

60 7

""U TUNACCEPTABLE 40

ACCEPTABLE

20

20 UNACCEPTABLE

(a) FOR PADDING T HICKNESS GREATER THAN ONE INCH

0 20 40 6o 8o 100

6AT[RUNACCEPTABLE ON

6 F40 ACC EPTABLE

cc,

UNACCEPTABLE

20

(b) FOR PADDING THICKNESS LESS THAN ONE INCH

0 V I I I I I I

0 20 40 60 80 100

STRAIN. PERCENTd

FIGURE 52. RECOMMENDED STRESSS-STRAIN PROPERTIES FOR PADDING

MATERIAL FOR HEAD CONTACT WITH CUSHION THICKNESS

OF AT LEAST 1.5 IN.

The impact surface is assumed to be flat; the data from which Figure 52 was

developed were obtained for simulated head impacts on flat surfaces with

energy levels up to 84 ft-lb, i.e., 11.2-lb head weight x 7.5-ft drop height.

The acceleration of the head should not exceed 120 G at an impact velocity of

20 ft/sec (or greater) while a higher level of acceleration can be sustained

at lower velocities (shorter pulse duration). This accounts for the

122
different stress-versus-strain values shown in Figure 52, i.e., a higher G or

crush stress is acceptable at the lower design velocity expected for the thin

padding.

The criteria of Figure 52 are to be satisfied by the padding material over

the entire anticipated operating temperature range if the potential for survival is to be maintained. However, practical considerations and risk analysis may reduce the temperature range requirements. Temperature sensitivity

must be considered as a padding material selection criterion. Other padding

material evaluation methods are discussed in Section 11.9.4 of Volume IV.

Stress-strain curves for several polyurethane-foamed plastics are shown in

Figure 53. The curve3 show that a density of 3 lb/ft3 or less will satisfy

the criteria of Figure 52 (superimposed as a crosshatched area) over at least

part of the operational temperature range.

5.13.10.5 Application of Padding Material. In the absence of data for

extremity impacts, it is assumed that padding material that is suitable for

head impact protection will be suitable also for protecting extremities.

Extremity impacts are not likely to have the potentially severe effects of

head impacts. It is suggested that areas within the extremity strike envelope having radii of 2 in. or less be padded and that such padding have a

minimum thickness of 0.75 in.

Caution must be exercised in padding sharp edges and corners. Padding installed in a manner that allows it to be broken away from the corner or cut

through by sharp edges offers no protection. It is recommended that edges

and corners to be padded have a minimum radius of 0.5 in. prior to padding.

A definite volume of the padding must be crushed to absorb the initial

kinetic energy of the head and protective helmet.

5.13.10.6 Ductile Materials. In cases where the use of padding material

is impractical or the thickness allowed is inadequate to provide the necessary protection, ductile energy-absorbing materials or frangible breakaway

panels should be used where possible. Window and door frames, control columns, electrical junction boxes, etc., should be designed with large radii

(1 in. or more) rather than with sharp edges and corners.

Swearingen concluded in Reference 65 that at impact velocities of 30 ft/sec

against rigid structure padded with materials even 6 in. thick, unconsciousness, concussion, and/or fatal head injuries will be produced. Where possible, a combination of deformable structure and padding material should be

considered to absorb the impact energy and to adequately distribute the

forces over the face. Surfaces to which this combination should be applied

are instrument panels, seat backs, bulkheads, and any other structure that

the head may impact during the crash sequence.

123
8011 I

70 :

60-

040

Lu

u ~40 -.

S301

/ LINE FOAM H.. TESTING OMASS 0 6PACT SP'ECIMEN

coo SIT 0'D IA ROr.0 VILOCITA (I.AD) 1zt

10 Ill'. ~ .'~g A 7.,STTC.S

*FOAM TAKFN FROM FOAM IN P LACE HELMET.

FOAM SFECWM " PITODUIJEO 1 OI" U.5 AIR FOISCE HELMET FOAMING KIh

WITH EQUAL rARTS Of POLYOL AND ISOCVAHATE.

0 04 0 8 0 100

COMPRESSIVE STRAIN, PERCENT

FIGURE 53. EFFECT OF DENSITY ON STRESS-STRAIN CURVES

FOR POLYURETHANE-FOAMED PLASTIC.

124
5.14 DESIGN CHECKLISTS

5.14.1 General Design Checklist

Yes No N/A

1. For load-linited seats, do all materials in critical

structural members possess a minimum elongation of

5 percent in the principal load direction?

2. For nonload-limited seats, do materials in critical

structural members possess a minimum of 10 percent

elongation?

3. Is there adherence to the flammability and toxicity

requirements of Chapter 6?

4. In load-limited portions of the seat, where loads

can be predicted accurately, are minimum margins of

safety for shear and tensile bolts 5 and 10 percent,

respectively?

5. In nonload-limited portions of the seat, are minimum

margins of safety for shear and tensile bolts 15 and

25 percent, respectively?

6. In the vicinity of welded joints, have crosssectional areas been increased by 10 percent to

account for uncertainties, stress

concentrations, etc.?

7. Have seat attachments been designed so that neither

buckling nor warping of the floor or bulkhead will

interfere with seat operation or seat integrity in

a crash?

8. Has the rcStraint s.. tem anchorane been designed son

that the restraint system will function effectively

as the seat strokes?

9. Is the use of castings avoided in the primary

seat structure?

10. If castings are used, are they sufficiently ductile,

or does the design allow for realistic seat deformation during crash load application without failure

of the castings?

11. Do nonmetallic materials comply with FAR 25?

12. Can troop seats be removed in 20 sec per occupant

position?

125 1
Yes No NiZA

5.14.2 Seat Strength and Deformation Checklist

1. Does the seat meet the longitudinal load-deformation

requirements of Figure 34?

2. Will the seat withstand a 12-G aftward load?

3. Is the vertical energy-absorption system designed

for a load factor of 14.5 G based on the effective

weight of the 50th-percentile aviator or trooper?- -

4. Does the crew seat possess a minimium vertical stroke

distance of 12 in. (from the lowest vertical adjustment position)?

5. Has the use of a variable-force energy absorber

been considered?

6. Does the troop seat possess a minimum of 17 in.

of vertical stroke?

7. Does the seat have a capability of withstanding an

upward load of 8 G?

8. Does the seat meet the lateral load-deformation

requirements of Figure 35?

9. Are the static attachment strengths for components

mounted on the seat, such as armored panels, based

on the following load factors:

0 Downward: 50 G

* Upward: 10 G

* Forward: 35 G

* Aftward: 15 G

e Lateral: 25 G

5.14.3 Seat Cushions Checklist

1. Are seat cushions of the type that minimize dynamic

overshoot in vertical deceleration?

2. Is the thizkness of the compressed seat cushion

between 0.5 and 0.75 in., or has it been demonstrated that the cushion design and material properties produce a beneficial result?

126
Yes No N'A

5.14.4 Litter Strength and Deformation Requirements Checkjist

1. Does the litter system possess the vertical strengthdeformation capability of Figure 41, based on an

occupant weight of 250 Ib?

2. Does the litter system possess the capability of

withstanding an upward load of 8 G?

3. Does the litter system meet the lateral loaddeformation requirements of Figure 42?

4. Can the litters be loaded laterally into the

aircraft?

5. Can the complete set of litters be loaded and

unloaded to flight readiness in 10 sec or less

in an emergency situation?

6. Does the litter system eliminate need for special

mounting hardware that remains attached to the

aircraft?

7. Can the standard cargo tiedown system be used as

the primary litter system attachment to the aircraft structure?

8. Will the litter installation accept the current

standard military litter?

9. Does the installation support the litter in such a

manner as to develop the maximum load-carrying

capability of the standard litter?

10. Would the litter installation be adaptable to a new

and improved military litter design?

11. Does the litter installation, when removed from the

aircraft, leave the aircraft free of all protuberances, brackets, and other objectionable

operational hazards?-

5.14.5 Restraint System Design Checklist

I. Are the lap belt anchor points located so that a

maximum angle of 55 degrees and a minimum angle of

45 degrees exists between the lap belt and the buttock reference line, as illustrated in Figure 29?

2. Is the point where the shoulder harness is attached

to or passes through the srat back between 26.5 and

27.5 in. above the seat reference point? --

127
Ye__•s B N/A

3. Does the shoulder harness anchorage or guide on the

seat back permit no more than 0.5-in, lateral

clearance?

4. Does the shoulder harness guide on the seat back

have a 0.25-in. minimum radius as illustrated in

Figure 32?

5. Is the lap belt tiedown strap (crotch strap) attached

to the seat pan centerline at a point 14 to 15 in.

forward of the seat back?

6. Are the forces required for adjustment of all webbing

item lengths no greater than 30 lb?

7. Are the lap belt adjusters located s, as to not

exert pressure on the iliac crests?

8. Are the shoulder strap adjusters located low enough

on the chest to avoid concentrated pressure on the

collar bones?

no 4^ Le restrain h" a...... sebsmblie moo the •

minimum load and maximum elongation requirements of

Table 8?

10. Have the stitched joints in the restraint harness

been designed according to the criteria discussed

in Section 5.7.4.3 and do the joints have a

30-percent marlin?

11. Is a minimum webbing thickness of 0.045 in. used on

all restraint harness components?

12. Do the restraint harness components meet the following minimum width requirements:

0 Lap Belt - 2.00 in.

* Shoulder strap - 2.00 in.

0 Tiedown Strap - 1.75 in.

13. Do oll webbing fittings, over which webbing is

wrapped, possess the 0.062-in. minimum radius

illustrated in Figure 32?

14. Does the restraint harness have a single-point

release system that can be released after being

exposed to design crash loads by exerting a 30-lb

force with one finger or a 50-lb force with on(

finger when supporting the entire weight of the

occupant?

128
Yes No NZA

15. Is the single-point release protected from inadvertent release?

5.14.6 Protective Padding Checklist

1. Are all areas within the extremity strike envelope,

having radii of 2 in. or less, padded with a minimuni

thickness of 0.75 in.?

2. Do padded corners of edges have a minimum unpadded

radius of 0.5 in.?

3. Are ductile energy-absorbing supports used where

possible under padding, particularly where head

impact is likely?

5.14.7 Cockpit Controls and Eguipment Checklist

1. Are rudder pedals separated from each other and from

adjacent structure by less than 2 in. or more than

6 in., as illustrated in Figure 50?

2. Are controls and control columns dcsigncd so that

fracture due to an occupant's striking the column

will occur at a point no more than 4 in. above the

pivot point, and so that the failure will be clean

without jagged or torn edges, or are they equipped

with an energy-absorbing section?

129
6. AIRCRAFT POSTCRASH SURVIVAL

6.1 INTRODUCTION

This chapter presents the criteria that are to be applied in designing post- crash survival into an aircraft. Although initial crash-resistance considerations, sL-h as maintaining structural integrity around the occupant and

reducing the crash forces transmitted to the occupant, are of primary importance in survival, hazardous postcrash conditions must be prevented or

reduced if the occupant is ultimately to survive. The threat of postcrash

fire must be minimized, and adequate escape and rescue provisions must be

incorporated into the aircraft.

This section includes criteria for designing fuel, oil, and hydraulic systems

to minimize the occurrence of postcrash fires; for selecting less flammable

interior materials; for selecting provisions that increase survival chances

during aircraft ditchings; and for designing emergency escape provisions and

crash locator beacons. The user is referred to Volume V for more complete

information and reference sources.

6.2 FUEL SYSTEM DESIGN CRITERIA

The following criteria are applicable to all auxiliary fuel systems, such as

ferry systems and extended range systems, as well as to the primary aircraft

fuel system.

6.2.1 General

The fuel system must be designed to minimize fuel spillage during and after

all survivable crash impacts. It must also be designed to prevent spillage

of fuel through the vents during a rollover or in any other adverse attitude. Spillage that cannot be avoided, such as that occurring during the

functioning of self-sealing breakaway couplings, must be precluded from ignition by controlling ignition sources (see Section 5.5 of Volume V).

6.2.2.1 Fuel Tank Location. The location of fuel tanks in an aircraft is

of considerable importance in minimizing the postcrash fire hazard. The

location must be considered with respect to occupants, ignition sources, and

probable impact a-eas. The fuel tanks should be located as far as possible

from probable impact areas and from areas where structural deformation might

cause crushing or penetration of the tank. If possiblc, fuel tanks should

not be installed:

• Immediately adjacent to occupiable areas.

* Immediately adjacent to engine compartments.

0 Immediately adjacent to electrical compartments.

* Under heavy masses, such as transmissions and engines.

• Near the bottom of the fuselage.

130
0 Over landing gears.

V In leading edges or anticipated failure areas of wings.

6.2.2.2 Fuel Tank Construction. Fuel tanks should have smooth, regular

shapes with the su4,,' area contoured gradually into the tank bottom. All concave corners should have a minimum radius of 3 in., and all convex corners a

minimum radius of 1 in.

All fuel tanks must be fabricated from crash-resistant material which meets

or exceeds the requirements of MIL-T-27422 (Reference 66). All fuel tank

fittings must have a tank pullout strength tha't meets or exceeds that specified in MIL-T-27422.

A self-sealing, breakaway, tank-to-tank coupling should be used wherever two

tanks are connected directly with no intervening fuel line.

6.2.3 Fuel Lines.

Fuel lines should be constructed and routed so as to withstand all survivable

crash impacts. This may be done by allowing the lines to elongate or shift

with deforming aircraft structure rather than being forced to carry high tensile loads.

rel L in o All fuel lines that could be readily

dai.iaged in an accident of severity up to that indicated in Iable 2 should

consist of flexible rubber hose with a steel-braided outer sheath, where

possible. The hoses should be capable of elongating 20 percent without the

hose assembly spilling any fuel. If "stretch 'le" (20-percent minimum

elongation) hoses are not used, all hoses should be 20 to 30 percent longer

than necessary to provide added length for structural displacement.

When the hose assemblies are subjected to Ipare tension loads or to loads applied at a 90-degree angle to the longitudinal axis of the end fitting, ;s

shown in Figure 54, hoses must not pull out of their end fittings, neither

should the end fittings break, at less than the minimum loads shown in

Table 17. Loads must be applied at a constant rate not exceeding 20 in./min.

The number of fuel line couplings should be held to a minimum. Wherever possible, a single, one-piece hose should be run through a bulkhead opening rather

than be attached to the bulkhead with rigid fittings. The opening should be

I in. larger in diameter than the hose diameter, with the hose stabilized by a

frangible panel or structure. A grommet should be installed in the opening to

preclude wear on the hose. Self-sealing breakaway couplings must be used whenever a line goes through a firewall so that the line will seal if the engine is

displaced during crash impdct. Breakaway couplings will not be required if the

engine is tied down to a strength level of 20 Gz, 20 GX, and 18 G , and

if the engine is located so that crushing of the lines and fittings is noL

likely in any survivable accident.

All fuel line-to-fuel tank connections should consist of self-sealieg breakaway couplings. These couplings must be recessed into the tank so that the

tank half does not protrude outside the tank wall more than 1/2 in. after

coupling separation. The shape of the tank coupling half should be basically

smooth to avoid snagging on adjacent structures or cutting the tank wall. An

131
Load Load

(4

Tension tests

Load Load

9 0-(leqrcc tests

FIGURE 54. HOSE ASSEMBLY TEST MODES.

acceptable substitute for a breakaway valve is a hose constructed of material

identical to thiat of the tank with an end fittir j strength equal to 80 percent of the tank tear-cut strength (MIL-T-27422, Paragraph 4.6.5).

132
TABLE 17. REQUIRED MINIMUM INDIVIDUAL LOADS FOR STANDARD

HOSE AND HOSE-END FITTING COMBINATIONS

Minimum Minimum

Hose End Fitting Tensile Load Bending Load

Fitting Type Size* (Ibj (lb)

STRAIGHT -4 575 450

Tension - -6 600 450

-•8 goo 700

-]0 1250 950

-12 1900 1050

-16 1950 1450

Bending = -20 2300 1600

-•24 2350 2750

-32 3500 4000

2. L LOW -4** 575 800

Tension - -6"1 600 850

-V** 900 !250

n-1 1250 575

-12 1900 675

-16 1950 1I0D

-20 2300 1250 •

Bending -24 2350 20?5

""32 3500 

3500

4_3_ -4** 575

Te-b** 600 425

-8** M00 425

J -10 1250 425

-12 1900 600

-10 1950 1000

-20 2300 1000

S g =-24 2350 2400

J-32 3500O 3700

*1 ittirj s1ize given in l1,I1 in. unts t i . -4 = 4/1i_

or 1/4 in

L lbow nna te r i a 1 s toel I

133
6.2.3.2 Fuel Line Location. Fuel lines should be located as far as possible from probable impact areas and areas where structural deformation can

cause crushing, penetration, or excessive tensile loading of the lines. When

fuel lines must be routed through areas of probable large displacement, such

as wing-to-fuselage attachment points, self-sealing breakaway couplings must

be incorporated into the lines to allow for complete line separation with a

minimum of fuel spillage.

Fuel lines should not be routed in the following areas:

* Near the bottonm of the fuselage.

* Over landing gears.

* Under, in front of, or at the sides of heavy masses,

such as engines and transmissions.

* In the leading edges of wings.

0 In anticipated areas of rotor blade impact.

* Adjacent to electrical wiring.

Fuel lines should not be routed through electrical compartments or occupiable

areas unless they are shrouded or otherwise designed to prevent spillage.

In order to protect the lines from impact damage, fuel lines should be routed

along heavier basic structural members wherever possible. All fuel lines

must be adequately supported by frangible clamps attached to other structure.

Fuel lines should be grouped together and exit a fuel tank in one centralized

location. This location should be in the area of the tank that. is least

vulnerable to anticipated crash loads and structural deformations. However,

ballistic vulnerability considerations may modify this requirement.

The number of fuel lines in the engine compartment should be minimized. When

mlore than one line enters an engine compartment, the lines should be grouped

together and pass through the firewall in a protected location unless the

structur I integrity of the firewall would be compromised.

6.2.4 Frangible Attachments

Frangible structures or frangible bolts should be used at all attachment

points between fuel tanks and aircraft structure to prevent fuel tank components from being torn out of the tank wall during impact. Frangible attachmerits should be used at other points in the flammable fluid systems where

aircraft structural deformation could lead to flammable fluid leakage.

The load required to separate a frangible attachment from its support struc--

ture must be between 25 and 50 percent of the minimum load required to fail

the weakest component in the attached system, as illustrated in Figure 55.

(The failure load of the attached system components may be determined either

by analytical computations or by testing methods based upon the failure modes

most likely to occur during crash impact.) To prevent inadvertent separation, failure loads should be at least five times normal operational and

service loads at the frangible attachment location.

134
Aircraft Tank wall

structure

Metal tank fitting Frangible bolt

Aircraft

structure

Flange Shear plane

ITEM LOWEST FAILUkE LOAD (LB)* iAILUvE MOvE

Aircraft

structure 4000 Shear

Tank fitting 3000 Pull out of

tank

Flange 5 00 Shear

Frangible bolt Not more than Break

3000 (tension-shear) --T-- = 1500

Not less than

300

--4--0 =750

*Loads may or may not be representative; values are for

explanatory purposes only.

FIGURE 55. SAMPLE FRANGIBLE ATTACHMENT SEPARATION LOAD CALCULATION.

A frangible attachment should separate whenever the required load (as defined

above) is applied in the modes most likely to occur during crash impact.

These modes--whether tension, shear, compression, or combinations thereof,

such as bending (tension-shear)--should be determined for each attachment by

analyzing the surrounding aircraft structure and probable impact forces and

directions.

135
All frangible devices should be statically tested in the three most likely

anticipated modes of separation. Test loads must be applied at a constant

rate, not exceeding 20 in./min, until failure occurs. In addition, all fran

gible attachments should be proof tested under dynamic loading conditions in

the three most likely anticipated modes of operation. The test load should

be applied in less than 0.005 sec, and the velocity change experienced by the

loading jig should be 36 + 3 ft/sec.

6.2.5 Self-Sealing Breakaway Valves

Self-sealing breakaway valves should be installed at all fuel-tank-to-fuelline connections, tank-to-tank interconnects, and at other points in the fuel

system where aircraft structural deformation could lead to system failure.

The valves should allow only a minimal amount of spillage upon separation and

should permit no external leakage when partially separated.

The load required to separate a breakaway valve should be between 25 and 50

percent of the minimum load required to fail the weakest component in the

attached system, as illustrated in Figure 56. To prevent inadvertent actuation during flight arid maintenance operations, the separation load must be

greater than five times normal operational and service loads at the coupling

location. To avoid complete or partial breakaway coupling separation during

maintenance operations, the separation load should never be less than 300 lb,

regardless of the fuel line size.

A breakaway valve should separate and seal whenever the required load (as

defined above) is applied in the modes most likely to occur during crash

impact. These modes, whether tension, shear, compression, or combinations

thereof, should be determined for each coupling by analyzing the surrounding

aircraft structure and probable impact forces and directions.

All breakaway valves should be subjected to static tensile and shear loads to

establish the load required for separation, nature of separation, leakage during valve actuation, general valve functioning, and leakage following valve

actuation. The rate of load application should nct be greater than

20 in./min. Tests to be used where applicable are shown in Figure 57.

in addition, all breakaway valves inust be proouf tested ui•wedY-r Udyna,,,,ic loald.in,

conditions. The valves must be tested in the three most likely anticipated

modes of separation. The test configurations should be similar to those

shown in Figure 57. The load should be applied in less than 0.005 sec, and

the velocity change experienced by the loading jig should be 36 ± 3 ft/sec.

All breakaway valves should incorporate positive provisions for ascertaining

that the valve is locked together during normal installation and serviLe. In

addition, all breakaway valves must incorporate provisions in their design to

prevent uncoupling due to operational shocks, vibrations, accelerations, etc.

6.2.6 Fuel Drains

All fuel tank drains should be recessed into the tank so that no part of the

drain protrudes outside the tank wall. All attachments of fuel drains to aircraft structure should be made with frangible fasteners.

136
AIRFRAME :STRUCTURE

HOSE END TANK

FLEX HOSE METAL TANK FITTING

BREAKAWAY VALVE

FRANGIBLE SECTION

tlowest Failure

Item Load (lb)* Failure Mode

Flex hose 3000 Tensile breakage

Flex hose 1500 Pull out of end fitting

Tank fitting 7500 Pull out of tank

Hose end coupling 1650 Break (bending)

Breakaway valve 2500 Pull out of tank fitting

Breakaway valve Not more than Break at frangible section

1 500 750 2

Not less than

1500 = 375

4

'Loads may or may not be representative, values are for

explanatory purposes only.

FIGURE 56. TYPICAL METHOD OF BREAKAWAY LOAD CALCULATION

FOR FUEL-TANK-TO-LINE BREAKAWAY VALVE.

The number of fuel line drains should be held to a minimum by designing the

fuel system to avoid low points in the lines. If drain lines are necessary,

they must be made of low-strength materials.

Drain valves for tanks and lines should be designed to be positive locking in

the closed position. Fuel drain actuation must not require the operator to

lie down under the aircraft. Drains should be located where discharged fuel

will not cause an added fire hazard.

137
Load Lad

Load

Static tensile test Static shear test

Load

Load

Hose

Load

Static bending Static shear test

(tensile-shear) test (tank-to-tank coupling)

FIGURE 57. STATIC TENSILE AND SHEAR TEST MODES

FOR SELF-SEALING BREAKAWAY VALVES.

6.2.7 Filler Units

The filler unit must be fastened to the structure with a frangible attachment,

and filler caps must be recessed into the tank wall to ensure that the cap

remains with the tank if the tank moves relative to the aircraft structure.

138
Long filler necks should be avoided if possible. If they must be used, they

should be fabricated from frangible materials and designed so that the filler

cap remains with the tank and does not snag on the aircraft structure during

impact.

Tank fillers must not be located adjacent to engine intakes or exhausts where

flaminable vapors could be ingested and ignited.

6.2.8 Fuel Boost Pumps

Boost pumps should be selected according to the following order of preference:

1. Suction system, engine-mounted pump.

2. Air-driven, tank-mounted or in-line pump.

3. In-line electric pump.

4. Electrically operated tank-mounted pump.

Pumps mounted within the fuel tanks should be rigidly bolted to the fuel tank

only. If the pump must be supported or attached to the aircraft structure, a

frangible attachment should he used.

The state of the art in fuel system design has shown that electrically driven

boost pumps can be eliminated. Air-driven boost pumns and engine-mounted

suction-type boost pumps now in oppration are much less hazardous alternative

solutions.

If electric boost pumps are used, the electrical wires must contain 6 in. of

extra length at the pump connection to accommodate crash-induced structural

deformation. The wires also must be shrouded to prevent their being cut during

crash impact. Nonsparking breakaway wire disconnects may be used in lieu of

the extra wire length.

6.2.9 Fuel Filters and Strainers

Fuel filters and strainers should not be located within the engine compartment

or adjacent to engine intakes or exhausts, if at all possible.

Filters and strainers should retain the smallest possible quantity of fuel.

Filters and strainers must have a structural attachment capable of withstanding

a 30-G load applied in any direction.

Self-sealing breakaway valves should be used to attach fuel lines to fuel filters and strainers in those locations where structural displacement is likely

to cause a separation of those components. Care should be taken to assure that

the valve, not the strainer or filter, is the weak link in the system.

139
6.2.10 Fuel Valves

The number of fuel valves should be kept to a minimum.

Large valves (e.g., fuel shutoff valves) should have a structural attachment

capable of withstanding a 30-G load applied in any direction. Self-sealing

breakaway couplings should be used at the valve-fuel line connections. Small

valves (e.g., check valves) must be fastened to the aircraft structure with

frangible attachments.

If electrically operated valves are used, they should be mounted on bulkheads

so that the electrical wires are on one side of the bulkhead and the valves

and lines are on the other side.

Section 4.3.3.9 of Volume V discusses the use of spillage control valves designed to stop the flow of fuel to the engine area when the engine is not

running, as in a crash.

6.2.11 Fuel Quantity Indicators

Fuel counters and float-type quantity indicators are preferred over rigid

capacitance probes to preclude puncture of the fuel tank during impact. If a

capacitance probe must be used, it should be fabricated from material processing as low a flexural rigidity as is consistent with operational requirements. A slightly rounded shoe should be incorporated at the probe bottom

end to avoid any tank-cutting tendency. Consideration should be given to the

use of frangible low-flexural rigidity curved probes to reduce the danger of

puncturing the tank during crash impact. The probe may also be mounted

frangibly or at an angle.

If tank-mounted quantity indicators must be attached to the aircraft structure, frangible attachments should be used.

6.2.12 Vents

Vent systems should be designed to prevent fuel flow through the vent lines

regardless of aircraft attitude or vent line failure. For this reason, high

strength fittings should be used between the metal insert in the tank and Lhe

vent line. If the vent outlet must be supported, it should be supported by

frangible attachments. The vent line should be made of wire-covered flexible

hose and should be routed so that it cannot be snagged in displacing structure during a crash. Self-sealing breakaway valves should be used at the

tank-to-lirie attachment if there is danger of the tank being torn free of the

supporting structure.

Vent lines should be routed inside the fuel tank in such a manner that if

rollover occurs spillage cannot continue. This can be accomplished with

siphon breaks and/or U-shaped traps in the line routing.

140
Antispillage vent valves inside the fuel tank are particularly advantageous

during rollover accidents and can be used in lieu of flexible lines, breakaway valves, and all other alternate considerations. These valves must be

designed and tested to demonstrate that:

0 The vent will remain fully open during all normal flight environmental conditions.

* The vent valves will close in extreme attitudes such as would occur

in a rollover.

* The vent valves will possess adequate venting capability under

critical icing conditions in flight.

If the fuel system is to be pressure refueled, a bypass system for tank overpressurization must be used. However, care must be taken to ensure that

spillage resulting from overpressurization due to tank compression during a

crash is released away from aircraft occupants and ignition sources.

6.3 OIL AND HYDRAULIC SYSTEM DESIGN CRITERIA

6.3.1 General

Even though oil and hydraulic fluids are carried in relatively small quantities, they are easily ignited and can serve, in turn, as ready ignition

sources for fuel. Therefore, oil and hydraulic fluid spillage should be prevented at all reasonable cost. The crash-resistance design criteria presented in Section 6.2 for fuel systems are generally applicable for oil and

hydraulic systems also.

6.3.2 011 and Hvdraulic Fluid Reservoirs

Oil tanks and hydraulic reservoirs should not be located where spilled or

sprayed fluid can readily be ingested into the engine or ignited by the

engine exhaust.

Oil +tans and hydrauili rsr,,rvinrs should not be located in the following

areas:

* Near the bottom of the fuselage.

* In or above engine compartments.

* In electrical compartments.

* In occupiable areas.

* Under, in front of, or at the side of heavy masses, such as engines

and transmissions, nor above landing gears.

Reservoir construction and mounting should be able to withstand 30-G forces

applied in any direction.

141
Oil tanks should be constructed from flexible, crash-resistant materials that

meet or exceed the strength and tear resistance required in MIL-T-27422 for

fuel tank material.

Alternatively, a metal tank can be used if it is in a relatively safe area

and is shielded and coated to prevent leakage in the event of a tank rupture.

6.3.3 Oil and Hydraulic Lines

6.3.3.1 Construction. Oil and hydraulic lines should consist of flexible

hoses with steel-braided outer sheaths, where possible. If the hoses cannot

elongate 20 percent without the hose assembly spilling fluid, 20 percent

extra length should be provided to compensate for structural displacement during a crash. All hose assemblies must meet the requirements of Table 16 when

tested as shown in Figure 54 (Section 62.3.1).

Where high-temperature operational requirements preclude the use of flexible

hose, coiled metal tubing should be used in areas where large crash deformation is expected.

The number of line couplings should be kept to a minimum. Wherever possible,

a single, one-piece hose should be routed through a bulkhead opening rather

than attached to the bulkhead with a rigid connection. The opening should be

1 in. larger in diameter than the hose diameter, with the hose stabilized by

a frangible panel or structure. However, self-sealing breakaway valves

should be used wherever a line goes through a firewall so that the line will

seal if the engine is displaced during crash impact.

Self-sealing breakaway valves should be used to connect flexible hoses to engines, oil tanks, hydraulic reservoirs, and system components if enough structural deformation to cause line elongation to the breakage point. iL probable.

When hydraulic or oil lines must be stabilized, they should be attached to

the aircraft structure with frangible fasteners.

6.3.3.2 Routing. Hydraulic or oil lines should not be routed in electrical oi occuplable areas unless they are shrouded to prevent, spillage.

,ydraulic or oil lines should not be routed in the following areas:

0 Near the bottom of the fuselage.

* Over landing gears.

I Under, in front of, or at the sides of heavy masses,

such as engines and transmissions.

a In the leading edges of wings.

* In areas of anticipated rotor blade impact.

* In any area where flaminable fluids could be spill2d or

sprayed onto hot surfaces or ingested into the engine.

* Above electrical wiring.

142
The number of hydraulic and oil lines in the engine compartme,,t should be kept

to a minimum. The lines should be grouped together and enter the engine compartment in a protected location.

6.3.4 Oi and Hydraulic System Components

System components (e.g., pumps, valves, filters, actuators) must not be located

in electrical compartments or occupiable areas. Components should not be located near the bottom of the fuselage or in the leading edges of the wings.

Components located in the engine compartment should be restricted to those ab--

solutely necessary for engine operation. For example, oil filters must not be

located here unless they are an integral part of the engine.

The construction and mounting of all system components must be able to withstand 30-G forces applied in any direction.

6.3.5 Oil Coolers

Oil coolers should not be located in the engine compartment, under the engine

or transmission, or in any area where oil could be spilled or sprayed onto hot

surfaces or ingested into the engine.

The oil cooler should be located as far as possible from anticipated impact

areas.

The oil cooler mounting(s) should be able to withstand 30-G forces applied in

any direction.

6.4 IGNITION SOURCE CONTROL CRITERIA

6.4.1 Llectrical Systems

6.4.1.1 Wiring. Electrical wires should be routed along heavier structural

members of the airframe wherever possible. Structural openings for wire

pass ag should be 8 to 12 tims !r, 1 r. n- d4 i ameter tha th wire. Sh- rp metal

edges should be protected by grommets to prevent chafing. Wire bundles should

be supported at frequent intervals along their length by frangible attachments

to the aircraft structure.

Wires that must pass through areas of anticipated structural deformation should

be approximately 20 to 30 percent longer than necessary. The extra length

should be accumulated in the form of loops or S-shaped patterns and located at

the areas of anticipated structural deformation.

""w- shoul b. routed abo-c cr aw.. from flam.mab. flud In, and they

should never be closely spaced between outer skin and fuel lines. Wires must

not be routed near flammable fluid tanks unless the wires are shrouded to prevent arcing. Wires should not be routed in the following areas:

* Near the bottom of the fuselage.

* Over landing gears.

* In the leading edges of wings.

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* In areas of anticipated rotor blade impacts.

* In areas of anticipated fuel spillage.

a Immediately adjacent to flammable fluid lines and vent openings.

Electrical wiring and components should be kept to a minimumr, in flammable

fluid tank areas.

Nonsparking breakaway connectors should be used in areas where excessive tensile loads may he applied, such as the wiiig-to-fuselage joint. All wire connectors must be of the shielded, nonsparking type.

6.4.1.2 Batteries and Electrical Accessories. Batteries and electrical

accessories should be, located as far as possible from flammable fluid tanks.

Batteries and accessories should be housed in compartments built into the airframe. These compartments sho-uld be lined with flexible, nonconductive, fireresistant panels ais sp~cified in Section 6.4.1.5.

Electrical wires shojuld exit the batteries and inverters on their least vulnerable side. The'o.- shoul'j be one full 6-in.-diameter loop of extra wire at the

battery and invcrter c.renections to accommodate crash-induced structural deformat ion.

The battery ,rid accessory mountings should withstand a force of 30 G applied

in any directiion.

6.4.1.3 Generators a.If Maqretos. If generators and magnetos are not engine mounted, thy should te installed in compartments built into the airframe. These cof.ipa.'tmeits should be located fairly high in the structure and

as far as po!siblo frcm, flammable fluids. The compartments shoild be lined

with panels ;.s specified in Section 6.4.1.5.

Electrical wires should exit the generators and magnetos on their least vulnerable side regardl.-ss of their location. The generator and magneto mountings

should wiLhstand 5 force of 30 G applied in any direction.

6.4.1.4 Liqhts and Antennas. Lights and antennas should be located as far

as possible from flammable fluids. Lights should be located as high as possible on the airframe structure. Landing lights should not be located in

front of wing fuel tanks.

lhe wires that attach to the lights should contain a 6-in.-diameter loop near A

the connection to accommodate crash-induced structural deformation.

6.4.1.5 Liners and Shrouds. Nonconductive paneling should be used as a

liner for all electrical compartments. The paneling materials should possess

a minimum tensile strength of 250 lb/in. of width and allow a minimum

elongation of 200 percent.

Monconductive material should be used to shroud all electrical wiring that

could be cut by deferming aircraft structure during crash impact. The shrouding materia' should meet or exceed a tensile load of 250 lb/in. of width and

should possess a minimum elongation capability of 200 percent.

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6.4.2 ShieldinM

Shielding should be used wherever necessary to prevent spilled flammable fluids

from reaching potential ignition sources or occupiable areas.

6.4.2.1 Spillage Barriers. Fuel tanks should be isolated from the occupants

by a minimum of two spillage barriers. These barriers may consist of the normal tank cavity chafing liner and the surrounding airframe structure. If the

chafing liner is considered as a barrier, it must be continuous structure completely encasing the fuel tank.

6.4.2.2 Firewalls. Firewalls should be designed to withstand all survivable

crash impacts without losing their structural integrity or sealing ability.

6.4.2.3 Fire Curtains. Fire curtains made from fire-resistant cloth may be

used to protect occupiable areas or ignition sources from flammable fluid spillage. Fire curtains may be installed in addition to, but not in place of, the

spillage barriers required in Section 6.4.2.1.

6.4.2.4 Flow Diverters. Drainage holes should be located in all flammable

fluid tank compartments to prevent the accumulation of spilled flammable fluids

within the aircraft. Drip fences arid/or drainage troughs should be used to

prevent the gravity flow of spilled fuels from reaching ignition sources, such

as hot engine areas or electrical compartments.

6.5 INTERIOR MATERIALS SELECTION CRITERIA

6.5.1 General

All aircraft interior materials such as seat fabrics and cushions, interior

wall insulations, and nonmetallic structural components must be flame resistant

and produce the least amount of smoke and toxic gases possible. Currently the

FAA flammability requirements specified in FAR 25.853 (Reference 67) are the

only specific mandatory requirements for aircraft interior materials. The FAA

amended the requirements in 1984 to add additional flammability tests for seat

cushions, except for crewmember seats, and is adding the Ohio State University

rate of heat release test procedure for interiur ceiling and wall panels, partitions, etc. Interior materials must be screened so that those with the least

smoke and toxicity emissions can be selected. Suitable screening tests are

referenced in Section 6.5.3.

6.5.2 FAR 25.853 Flammability Requirements

Materials used in each compartment occupied by the crew or passengers must meet

the following requirements:

0 Ceiling panels, wall Danels, partitions, structural flooring, etc.

Must be self-extinguishing when tested vertically by applying a

1550 OF flame to the lower edge of the specimen for 60 sec. Average

burn length not to exceed 6 in.; average flame time after removal of

test flame not to exceed 15 sec. Drippings may not continue to flame

more than an average of 3 sec. In addition, materials must meet the

OSU heat release rate in a vertical position exposed to a total heat

flux on the specimen of 3.5 watts per square centimeter. The average

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total heat release must not exceed 65 kilowatt-minutes per square

meter, and the alverage peak heat release rate must not exceed

65 kilowatts per square meter.

Floor coverings, textiles (including upholstery}, seat cushions.

paddingS, insulations (except electrical insulation), etc. Must be

self-extinguishing when tested vertically by applying a 1550 OF

flame to th2 lower edge of the specimen for 12 sec. Average burn

length riot to exceed 8 in., average flame time after removal of test

flame not to exceed 15 sec. Drippings may not continue to flame

more than an average of 5 sec. In addition, seat cushions must

meet an oil burner test. This test exposes the side of the seat

cushion to a specified oil burner for 2 min. During the next 5 min.

the burn length must not reach the side of the cushion opposite the

burner and must not exceed 17 in. Also, the average percentage

weight loss must not exceed 10 percent.

* Acrylic windgows_,_signs, restraint systems, etc. May not have an

average burn rate greater than 2.5 in./min when tested horizontally

by applying a 1550 OF flame to the specimen edge for 15 sec.

See References 67 and 68 for the complete text of the regulations and test

requirements.

6.5.3 Smoke and Toxic Gas Test Criteria

The FAA has not adopted criteria for smoke or toxic gas emissions from interior materials because the full-scale fire tests have demonstrated a correlation between flammability and smoke emission characteristics of the materials

tcsted. Also, the full-scale tests showed that there was a significant correlation between flammability and toxic emissions and that severe hazard from

toxic emissions does not occur until a flashover occurs. In addition, there

has not been good correlation shown between any of the laboratory tests for

smoke and toxic gases and full-scale fire tests. It should be emphasized,

however, that these generalizations are true only for the materials that have

so far been tested in the full-scale tests and only for the full-scale tests

SiMulating a fuel fire uLSide of the fuselage. It is possible that* n the

future, after more work has been done on the laboratory tests, some criteria

may be adopted.

In the meantime, screening tests should certainly be conducted on candidate

materials and systems to enable the designer to select those materials with

the lowest smoke and toxicity emissions arid to preclude using materials which

might generate high levels of smoke and toxic gases. It is recommended that

materials be screened for smoke emissions using either the test procedure for

the OSU release rate apparatus specified by NFPA 263 (Reference 69) or the

modified NBS smoke chamber as outlined in Reference 70.

The screening method to distinguish materials producing more toxic combustion

products than those from other materials should be performed using the NBS

toxicity test method (Reference 71). In this test, one material is considered significantly more toxic than another material if the toxic concentrations generated differ by an order of magnitude.

146
If fire-retardant coatings are used for fabric and trim materials, the effects, if any, of routine maintenance and cleaning procedures must be assessed. If the coatings can be removed by routine cleaning procedures, the

flammability and smoke/toxic fume tests should be repeated after a representative number of cleaning cycles.

6.6 DITCHING CRITERIA

6.6.1 General

Occupant survival during a ditching is highly dependent on egressing rapidly

from the aircraft before it sinks. This is especially true in helicopters, which tend to rol inverted and sink very rapidly. Disorientation and poor

underwater visibility further hamper successful egress. Available escape

times from helicopters range from a few seconds to a few minutes. The availability of emergency exits, adequate emergency exit lighting, and helicopter

flotation provisions can all increase the available escape time. Adequate

and easily deployed ditching equipment increases the probability of survival

after successful egress.

6.6.2 Emergency Exits

All U.S. Army aircraft should meet the criteria for emergency exits contained

in Section 6.7. Passenger-carrying helicopters operating over water environ- ments, however, shou d- 21 f II1 U - - .1 . ... ... . 4.. -L , L - - r.

normally be provided. Additional escape exits should be provided in the over- head, deck, and tail sections.

Explosively created exit systems should be considered because of their rapid

initiation times and immunity to the crash environment. Linear-shaped charges

should be placed around and extend beyond existing windows and hatches to

preclude the problem of jammed or stuck exits. Strategically placed shaped

charges in the overhead, deck, empty bulkhead spaces, etc,, can provide the

additional emergency exits required in the ditching environment. Criteria for

these types of systems are contained in Section 6.7.

6.6.3 Underwater Emergency LightinS

Emergency exits should be lighted with high intensity lights if they are to

be seen underwater. The required brightness of the lights depends on the

turbidity of the water, the distance between the observer and the light, and

the threshold sensitivity of the observer's eyes.

The escape hatch lights should have the highest brightness level of light permitted by other design conditions (up to 200 fL). Light tubes configured in

an inverted U around each hatch are most effective for outlining escape

hatches.

6.6.4 Helicopter Flotation Systems

An adequate number of helicopter flotation devices should be provided. Combinations of flotation methods, such as sponsons in conjunction with flotation

bags, sealed hulls, etc., should be used.

147
Sponsons can help stabilize a helicopter in relatively calm seas. However,

they must be quite large to be of any value in providing flotation to counteract the inherent instability due to a helicopter's high center of gravity.

Calculated aircraft stability must be verified by data from tests performed

on the aircraft or on a scaled model thereof.

The calculated stability afforded by flotation bags also must be verified by

test data. To achieve maximum effectiveness, the bags must inflate simultaneously prior to or upon water contact at slow speeds. Reliability of a

flotation bag system is of prime importance.

6.6.5 Ditching Equipment

Tie-down or stowage locations must be provided for life rafts, life preservers, survival kits, and miscellaneous ditching equipment, Restraint devices

and supporting structures must be designed to restrain the equipment to

static loads of 50 G downward, 10 G upward, 35 G forward, 15 G aftward, and

25 G sideward. All survival equipment must be readily available and easily

released from restraining devices after ditching.

Life raft mountings and restraining devices must be located and designed so

that rafts can be removed and deployed outside the aircraft within 30 sec

from the time the release or removal action is initiated.

When exterior installations for life rafts or other survival equipment are

provided, the mountings and restraining devices must be recoverable from an

exit intended for use in ditching. Release mechanisms must be designed to

minimize the possibility of jamming due to structural deformation incurred

during ditching.

6.7 EMERGENCY ESCAPE AND RESCUE DESIGN CRITERIA

6.7.1 Emergency Exits

6.7.1.1 General. Exits of sufficient size and number must be provided to

ensure that all occupants can evacuate the aircraft before postcrash conditions become intolerable, even if half of the exits are blocked. if a crashresistant fuel system is not installed, the maximum number of personnel to be

carried must be able to evacuate the aircraft within 10 sec. The allowable

evacuation time can be extended to 30 sec if a crash-resistant fuel system is

installed in the aircraft. The emergency exit criteria presented in this

chapter are predicated on a 30-sec evacuation time.

6.7.1.2 Types of Exits. A Class C exit constitutes tl- ninimum requirement for an emergency exit. (A Class C exit is a winouw, door, hatch, or

other exit intended primarily for emergency evaruation). Class C exit closures must be capable of being removed from the exit opening within 5 sec

regardless of the aircraft's attitude.

A Class B exit consists of a door, hatch, or other exit intended primarily

for service or logistic purposes (e.g., cargo hatches and rear loading ramps

or clamshell doors). Class B exits may be used instead of Class C exits if

adequate emergency releases are installed. A Class A exit (doors, hatches,

148
etc., intended primarily for normal entry and exit) generally may be used in

lieu of a Class C exit; however, if either Class B or Class A openings are

used in place of Class C exits, they must meet the 5-sec opening requirement.

6.7.1.3 Size of Exits. All exits must be sufficient in size and shape

to allow 95th-percentile combat-equipped troops and aviators to pass through

the exit at a rate of 1.5 sec per man or less. Therefore, Class C exits must

be a minimum of 22 in. in diameter, or 22 in. square, with 6-in. radius

corners, although larger exits are recommended. Other shapes may be used if

the minimum dimensions are met or exceeded.

6.7.1.4 Number of Exits. Each flight crew member must have access to at

least one usable emergency exit regardless of the attitude of the aircraft

after impact. When sliding or clamshell canopies are used, Class C exits

must be provided for crew escape in case the postimpact attitude of the

aircraft prevents jettisoning of the canopy.

A minimum of two Class C exits (or equivalent) must be provided in troop/

passenger sections, one on each side of the fuselage. Cockpit exits may not

be counted toward this requirement. Additional exits must be provided whenever the ratio of seats to passengers exceeds the 1-to-l0 ratio (e.g., if the

capacity is 21, three exits are required). These requirements also apply to

cargo compartments if the compartments have a capability for troop transport.

6.7.1.5 Location ofExits. Emergency exits must be equally divided between both sides of the aircraft to provide alternate means of escape if, for

any reason, the exits on one side become blocked. If feasible, in order to

prevent crowdiny during evacuation, side exits should not be located directly

across from each other. At least one exit on each side must be well above

the anticipated waterline during a ditching.

If the width of the fuselage between side exits is 5 ft or more, at least one

additional Class C exit should be provided overhead so that easy access to an

exit is available when the aircraft comes to rest on its side. If more than

20 occupants can occupy the troop/passenger section, one overhead exit should

be provided for every 20 occupants. If overhead exits are not feasible, bottom or fore and• or a ' - ex s may e provided instead. Alternati vy, siE

exits may be located where interior aircraft structures or components can be

used as steps to gain access to the upside exits. Such component-steps

should be able to support at least 300 lb. They should also n.aintain their

structural integrity and attachment to the aircraft when exposed to static

loads of 50 G downward, 10 G upward, 35 G forward, 15 G aftward, and 25 G

sideward.

Emergency exits should not be located in the following areas:

a In close proximity to the main landing gear.

* Under heavy components, such as engines and transmissions.

* In any area where it is necessary to move equipment, cargo, etc.,

to gain access to the exit.

149
a In any area where external components, such as engines or armament,

will interfere with occupant escape.

* Near potential fuel spillage areas.

9 Near major ignition sources, such as hot engines.

6.7.1.6 Operation-of Exits. The method of releasing and opening an emergency exit should be simple, obvious, and natural to all personnel carried in

the aircraft. All emergency exits should be capable of being completely

opened within 5 sec after the person initiating the action first places his

hand on the release handle.

Exit release mechanisms should permit release handle actuation and exit opening by one person using one hand. The releasing action should be natural to

the position of the operator initiating the actien and should be a continuous

motion from start to finish without sharp changes in direction. Secondary

operations should not be necessary. The final motion of the release handle

should contribute to the opening of the exit.

Release handles must be located on the exit closures themselves, or immediately adjacent to the exit openings, so that they are readily accessible.

However, the handles should not obstruct the removal of the exit closure or

impede escape through the exit opening. Release handles in cockpits and

troop compartments should be located so that crew members need not unlock

their shoulder harnesses in order to actuate the release mechanism.

Accidental release of exits in flight should be prevented. Release mechanisms should be designed so that improper or incomplete closing of the exit

closure will be obvious. Easily removable protective covers may be used to

prevent inadvertent actuation of exit release handles.

It is essential that all emergency exits be designed so that rescue personnel

can open them from outside the aircraft. Internal and external release mechanisris should be design2d so that they can be actuated simultaneously without

interfering with each other. Means to prevent icing of the outside release

mechanisms and handle mounts should be provided.

Once the release mechanism has been actuated, only the single operation of

pulling or pushing the exit closure into the clear should be necessary. All

emergency exit closures should be designed to fall free or be easily pushed

outward if the aircraft is not pressurized. In pressurized aircraft, exit

closures should be removed inwardly, but if possible they should then be

canted at an angle and pushed out the exit opening. "Push out"-type Class C

exits should also be designed so that they can be pushed in from the outside

by rescue personnel.

Emergency ex;ts should be designed to permit removal of the exit closure in

spite of seal vulcanization, ice accumulation, and moderate fuselage deformation. A peripheral clearance of at least 0.20 in., provided between the exit

closure and its frame, will help accomplish this goal.

6.7.1.7 Explosively Created Exits. Explosive systems for cutting emergency exits through existing doors and windows and through fuselage structures should be considered. These systems provide the advantages of

150
extremely rapid release times, simplicity of operation, and immunity to

jamming. If an explosive exit system is incorporated in the aircraft, the

following design criteria apply.

The arming/firing system should be designed for simple and rapid actuation of

the explosive system, yet provide maximum safety against inadvertent actuation. Arming and firing should be accomplished in two separate and deliberate actions, with the arming function always under the control of the

flight crew. The safe/arm mechanism should remain in its chosen position

(armed or disarmed) until a deliberate action to change its position is

initiated. The safe/arm mechanism should not change positions due to system

failure, or due to any environmental or crash inputs. Disarming capability

must be provided to permit safing the system when normal safing modes are

inoperable.

The firing mechanism should be independent of any external energy source.

Firing mechanisms should be located adjacent to each emergency exit so that

each exit can be opened independently, From both inside and outside the

aircraft.

The linear shaped charges used to cut the exit openings should be held

securely in position against the aircraft structure. The size of the exit

openings should conform to Class C requirements. The jettisonable section

should be ejected outward. Energy-absorbing backup material should be placed behind the shaped charge to control the backblast of the explosive.

All explosives used in the system should possess as high a thermal limit as

possible. The system should be able to function when exposed to ambient air

temperatures up to 40 OF, yet not function during brief exposure (30 to

60 sec) to postcrash fires. The system should be designed to minimize the

possibility of system actuation igniting any spilled fuel. Thus, the amount

and duration of any exposed flame should be minimal.

6.7.1.8 Access to Exits. Access from aisles to all exits should be provided so that exits are not obstructed by any aircraft structures or components that would impede escape. The width of aisles at any point between

seat rows must allow unobstructed movement of 95th-percentile troops with

full combat equipment. Therefore, the aisle width should be at least 17 in.

Where it is necessary to pass through seat rows to gain access to emergency exits, the longitudinal spacing between the rows should be sufficient to

permit these troops to move at a rate consistent with the capacity of the

exit (1.5 sec per man or less).

6.7.2 Emergency Lighting

6.7.2.1 Interior Emergency Lighting. Interior emergency lighting should

provide sufficient illumination throughout cockpit and cabin areas to permit

occupants to locate emergency exits and survival equipment, perceive escape

paths, and avoid obstacles while moving toward the exits. Minimum average

illumination in clear air along passageways leading to each exit and in front

of each exit should be 0.05 fc measured 20 in. above the floor (excluding

canopy aircraft).

151
6.7.2.2 Emergency Exit Lights. Supplementary emergency lighting units,

with adequate brightness to permit occupants to identify exits, read exit

operating instructions, and actuate exit release mechanisms during reduced

visibility conditions (darkness, smoke, etc.), should be provided at or near

each emergency exit. Exit lights should be mounted in the lower part of the

cabin to the extent possible. All passenger/troop-carrying aircraft must contain internally illuminated exit sign, with a minimum brightness of at least

25 fL, although brighter lights are strongly recommended. Aircraft whose mission requirements include troop transport over water should contain exit sign

lighting that meets the requirements specified in Section 6.6.3. Canopy

aircraft may be excluded from these requirements.

6.7.2.3 Exterior Emergency Lighti~ng. For noncombat missions, exterior

emergency lighting should be considered to iiluminate the ground near each

exit and in areas where escape and survival equipment will be deployed. The

light intensity on the ground should be 0.02 fc minimum.

6.7.2.4 Structural Requirements. All emergency lighting units should be

self-contained, explosion-proof, operable under water, and accessible for

periodic maintenance. To ensure structural integrity and continued operation

after a crash, the lighting system should be capable of withstanding the following crash loads: 50 G downward, 10 G upward, 35 G forward, 15 G aftward,

25 G lateral. The crash environment is more fully defined by the velocity

changes presented in Table 2. Except for those lights directly destroyed by

the crash, breakup of the fuselage should not render any portion of the

lighting system inoperative.

6.7.2.5 Power Sources. All units should be capable of operating independently of thj. main aircraft lighting system. Emergency lighting power

sources should be independent of the main power source of the aircraft. They

should contain power sufficient to provide effective illumination for a

minimum of 15 min.

6.7.2.6 Actuation of Lighting Units. Emergency lighting units should be

dCtuated automatically in as many survivable accidents as possible. This can

be accomplished by using inertia sensors capable of sensing lower-severity

accidents. Sensor criteria should be identical to those specified for crash

locator beacons in Section 6.8. An override switch to nullify the automatic

feature when desired must be provided. Manual actuating switches should be

provided so that emergency lights can be turned on prior to a crash if

desirable.

6.7.3 Emgrency_ Exit Markings

lmr(rgýnicy cyits 0,ould be clearly marked both inside and outside the aircraft. I, addition, instructions for releasing the exits should be clearly

marked besidc the exit release mechanisms.

All U.S. Army aircraft must be painted ard marked ,.ccording to the requiremtrits of lb 746-93-2 (Reference 72). Although these requirements are summariLJ in Volume V of this guide, the reader is referred to TB 746-93-2 for comiilet'e detail b.

1:2
6.7.4 Crew Chief Stations

At least one crew chief station should be located in each troop compartment.

The station should be located as near the main or emergency exits as possible

and should provide complete surveillance of the troop compartment.

6.7.5 A) •rm Systems

Aircraft with passenger or troop compartments should be equipped with an audible emergency alarm device that can be heard over the highest decibel noise

level expected in the aircraft. Consideration should be given to providing

visual as well as audible warnings.

6.8 CRASH LOCATOR BEACON DESIGN CRITERIA

6.8.1 General

Crash locator beacons may be fixed, portable, or deployable, as specified by

the procuring activity according to its aircraft mission requirements.

Fixed equipment is permanently mounted in the aircraft. Although the transmitter, antenna, and power supply need riot be contained in one package, their

close proximity to each other will reduce the chances of connecting circuitry

being damaged during crash impact.

Portable and automafically deployed beacons should contain the transmitter,

anterna, and power supply in one package. Portable beacons must be easily

removed from their installations by crew members, yet their installations

must be secure enough to protect them from impact damage.

Automatically deployed beacons should be designed to withstand ground impact

forces following their ejection. They should also be buoyant, self-righting,

and stable when floating in water and not adversely affected by immersion in

fresh or salt water for the life of the power supply.

Crash locator beacons may be either manually or automatically activated.

Sincc automatic activation requires no previous action on the part of the

crew, it is the preferred method. However, an arming switch should be provided so that automatic activation can be used or not, depending on the air

craft mission. A manual activation switch also should be provided so that

the beacon can be activated if the arming switch is not on or if, for any

other reason, the beacon is not automatically activated.

6.8.2 C(ash Sensors

Although different types of crash sensors might be used, the current state of

the art is such that inertia sensors are the preferred choice. Systems have

been designed which include self-testing diagnostic features. Regardless of

the type of sensor used, the sensor must be responsive to the majority of

survivable aircraft accidents, including those accidents in which the crash

forces and damage are minimal. At the same time, the sensor should ignore

normal vibrational loads and flight loads up to the limits of

maneuverability.

153
In order to sense 75 tu 80 percent of light fixed-wing accidents, an inertia

sensor should have a sensing threshold of 2 G. Although the 2-G threshold

level is below the accelerations sometimes experienced during flight, the

inertia sensor can be designed to filter out vibration and flight loads if it

also must detect a velocity change typical of crash rather than operational

conditions before it actuates.

Since most fixed-wing aircraft accidents have a major longitudinal component of velocity arid force, a unidirectional inertia sensor mounted with the

active axis forward in the direction of the longitudinal axis of the aircraft

is sufficient. A longitudinal inertia sensor should be designed to actuate

at a threshold of 2-G acceleration and a minimum velocity change of

3.5 ft/sec.

The above specifications are also satisfactury for rotary-wing aircraft in

the longitudinal direction. However, since helicopters often have large

vertical crash forces with minimal longitudinal forces, a vertically oriented

crash sinsor should be employed in addition to a longitudinal sensor. The

vertical sensor should be designed to actuate at a higher acceleration level

and velocity change. Specific levels for helicopter operations have been

established in Reference 73.

The sensor should be able to withstand impact forces associated with severe

survivable crashes arid still function. Thus the sensor should withstand

shock pulses equal to or greater than those required for the transmitter,

power Sup l , atiU ailLteiia.

The inertia sensor criteria should be based on crash forces typical of those

experienced in the occupant compartment during survivable crashes. Therefore, the sensor should be located in an area that will experience crash

forces representative of those in the occupant compartment. The sensor

should, of course, be protected from possible impact damage.

The sensor should be mounted to rigid structure to prevent the amplification

or attenuation of flight or crash loads that can occur with flexible structures. For the same reason, soft mounting materials, such as flexible straps

or Velcro fasteners, must not be used.

6.8.3 Transmitters

Operating frequencies and transmitter signal characteristics and power should

be determined by the procuring activity according to its own needs and the

available detector systems. A

The transmitter should be designed so that it can be either manually or automatically activated. An arm switch should be provided so that automatic activation can be selected or not, as desired.

A cockpit warning light or sound should be provided to alert the crew to inadvertent transmitter activation.

The transmitter should be located in an area that is least likely to be subjected to impact damage. The transmitter and its mounting should be designed

to withstand the impact forceý of a severe survivable accident without comproroising the operation of the transmitter.

154
6.8.4 Antennas

The antennas, except for those used in portable and deployable beacons, are

usually mounted outside the aircraft. The antennas should be located away

from anticipated impact areas, such as the front or bottom of the aircraft,

wing or tail surfaces likely to impact trees, etc., and those portions of

helicopters apt to experience rotor blade strikes during impact. The antenna

mounting should be able to withstand the decelerative forces of severe survivable impacts. Low profile antennas have been developed For the normal emergency frequencies.

6.8.5 Power Supplies

The crash locator beacon should have its own independent power supply so that

it is not dependent on aircraft power for its operation. The power supply

should be capable of providing the necessary power for optimum transmitter

operation over the time period and under the environmental conditions specified for the particular aircraft.

If the power supply is not integral with the transmitter, it should be

mounted to the aircraft in a location away from anticipated impact areas and

should have an attachment strength equal to that of the transmitter.

All electrical wiring between components of the system should be protected

from impact damage unless the components are packaged together. Protection

can be accomplished by following the criteria in Section 6.4.i.i.

6.9 DESIGN CHECKLISTS

6.9.1 Fuel System Design Checklist

6.9.1.1 Fuel Tanks

Yes No N/A

1. Are the fuel tanks located as far as possible from

anticipated impact areas, occupiable areas, large

weight masses, and nrimary ignition sources?

2. Are the fuel tanks located as high up in the

structure as possible?

3. Are the fuel tanks located where there is no danger

of puncture by a collapsing landing gear?

4. Are the fuel tanks located so that transmissions,

engines, and similar massive componerits will not

crush the tanks during a crash?

5. Are the fuel tanks relatively safe from peietrative

damage by structural str-ingers and stiffeners?

155
Yes No N/A

6. Can each fuel tank displace in the airframe structure without tearing or inducing leaks around the

filler area, the fuel line entry and exit, the

quantity indicator, and the tank-to-structure

attachment points?

7. Do the fuel tanks have smcuth, regular shapes, with the sump gradually contoured into the tank bottom?

8. Do all fuel tank concave corners have a minimum

radius of 3 in., and all convex corners a minimum

radius of 1 in.? -

9. Do all fucl tanks mnet or exceed the requirements

of MIL-T-27422?

10. Do all fuel tank fittings meet or exceed the tank

pullout strength specified in MIL-T-27422?

6.9-1.2 Fuel Lines

11. Are all fuel lines made from flexible hose with a

steel-braided outer sheath?

12. Do all hose assemblies meet the strength requirements listed in Table 17, Section 6.2.3.1?

13. Can all hoses elongate 20 percent without the hose

assemblies spilling fuel?

14. Do fuel lines exit the fuel tank in one protected

location?

15. Has the number of fuel lines in the engine compartment been keot to a minimum?

16. Are fuel lines routed along heavier structural 4

members wherever possible?

17. Is as much of the fuel line as possible routed

through the fuel tanks?

18. Are fuel lines routed as far as possible from occupiable areas and electrical compartments?

19. Are fuel lines routed as far as possible from all

electrical equipment and wires?

20. Are fuel lines routed away from areas where large

structural damage is likely durtig a crash?

21. Are fuel lines routed away from the exhaust system

and hinh-temperature heating ducts?

156
Yes No NA

22. Are the fuel system lines designed with as few

fittings as possible?

23. Are the fuel system lines designed so that uncut

hoses are run through bulkheads rather than attached

to the bulkheads with fittings?

24. Are self-sealing breakaway valves used wherever a

fuel line goes through a firewall or bulkhead or is

attached to the bulkhead?

25. Are lines entering and exiting in-line boost pumps

made of flexible hose that is approximately 20 percent longer than necessary?

26. If fuel lines are not longer than necessary for

in-line boost pumps, are self-sealing breakaway

valves used in the lines near the boost pump!?

27. Are self-sealing breakaway valves used at all

points in the fuel lines where aircraft structural deformation could lead to line failure?

28. Are fuel line supports fraible to ensur-e release

of the line from the structure during crash impact? -

29. Will the frangible supports meet all operationalI

and service loads of the aircraft?

30. Are all continuous lines running through bulkheads

stabilized by frangible panels?

6.9.1.3 Frangible Attachments

31. Are frangible attachments used at all attachment

points betweev; thq fuel tanks and aircraft structure?

32. Do the specified frangible tank attachment separation

loads exceed all operational and service loads by a

satisfactory margin?

33. Are the specificd frangible attachment separation

loads between 25 and 50 percent of the loads required to fail the attached system or components? m

34. Will the frangible attachments separate whenever the

required loads are applied in all possible modes

likely to occur during crash impacts? -

1

1157 m
Yes No N/A

6.9,1.4 Self-Sealing Breakaway Valves

35. Are breakaway valves installed in all fuel-tank-tofuel-line connections, tank-to-tank interconnects,

and at other points in the fuel system where aircraft structural deformation could lead to system

failure?

36. Are the shapes of the breakaway valves remaining

in the fuel tank basically smooth?

37. Are the breakaway valves recessed into the tank

wall so that the tank half does not protrude outside the tank wall more than 1/2 in. after valve

separation?

38. Do the specified breakaway valve separation loads

exceed all operational and service loads of the

aircraft?

39. Are the specified breakaway valve separdtion loads

between 25 and 50 percent of the loads required to

fail the attached components or lines?

40. Atre ",,e L ',awa valves -e.Ci " i vered tn seVprAte

whenever the required *ioads are applied in thc

modes most likely to occur during crash impacts? -

6.9.1.5 Fuel Drains

41. Are all fuel line diain valves stabilized where

necessary with frangible attachments?

42. Are all structural attachments of fuel tank

drains made with frangible attachments?

43. Are all fuel tank drains recessed into the tank

so that no part of the drain protrudes outside the

tank wall?

6.9.1.6 Filler Units

44. Are filler units attached to the aircraft structure

with frangible attachments?

45. Are filler caps recessed into the fuel tank wall?

46. Are long filler necks avoided?

47. If filler necks are used, are they made from frangible materials and designed so that the filler cap

stays with the tank after filler neck separation?

158
Yes No N/A

6.9.1.7 Boost Pumps

48. Can an engine-mounted, engine-driven boost pump

be used in the aircraft?

49. If an engine-mounted suction system cannot be

used, can an air-driven boost pump be used?

50. Do in-line boost pumps have a structural attachment capable of withstanding a 30-G load applied

in any direction?

51. Are tank-mounted boost pumps fastened to the

structure with frangible attachments?

4 6.9.1.8 Fuel Filters and Strainers

52. Are fuel filters and strainers mounted outside the

engine compartment wherever possible?

53. Do all strainers and filters have a structural

attachment capable of withstanding a 30-G load

applied in any direction?

54. Do all strainers and filters retain as small a

quantity of fuel as possible?

6.9.1.9 Fuel Valves

55. Has the number of fuel valves been kept to

the minimum required for operation?

56. Are self-sealing breakaway valves used at all

valve-to-fuel-line connections where crash-induced

line failure is likely?

57. Are all small in-line valves fastened to the

structure with frangible attachments?

58. Do large valves have a structural attachment capable of withstanding 30-G loads in any direction?

59. Are fuel shut-off valves located outside the engine

compartment, either on the outside face of the firewall or at the fuel tank outlets?

6.9.1.10 Fuel Quantity Indicators

60. Can float-type quantity indicators be used in this

fuel system?

159
Yes No N/A

61. If probe-type indicators are used, are they fabricated from material that either is frangible or

possesses as low a flexural rigidity as possible?

62. Is a slightly rounded shoe incorporated at the

probe bottom end of all probe-type indicators, or

is the probe mounted at an angle toward the rear

of the aircraft?

63. Are frangible attachments used where it is necessary to stabilize the indicator by fastening

it to the structure?

6.9.1.11 Vent Systems

64. Are high-strength fittings used between the metal

insert in the tank and the vent line?

65. If vent outlets must be supported, are they supported by frangible attachments to the structure?

66. Is the vent line made of wire-covered flexible hose?

67. Is the vent line routed so that it cannot be

snagged in displacing structure during a crash?

68. Is a self-sealing breakaway valve used at the tankto-line attachment if there is danger of the tank

being torn free of the supporting structure?

69. Are vent lines routed inside the fuel tank in such

a manner that spillage cannot continue after a rollover accident?

70. if ai antispilIlage vent valve is used inside the _ tank in lieu of the above items, will the valve remain fully open during all normal flight conditions?

71. Will the vent valve close in the extreme attitudes

that will occur during a rollover?

72. Will the vent valve possess adequate venting cap

ability under critical icing conditions in flight?

73. If the fuel system is to be pressure refueled, is a

bypass system provided in case of tank overpressurization?

74. Is any spillage due to tank overpressurization

released away from aircraft occupants and ignition

sources?

160
Ye No_ N/A

6.M.2 Oil and Hydraulic System Design Checklist

6.9.2.1 Oil Tanks and Hydraulic Reservoirs

1. Are the tanks and reservoirs located as far as

possible from anticipated impact areas, occupiable

areas, large weight masses, and primary ignition

sources?

2. Are the tanks and reservoirs located as high up in

the structure as possible?

3. Are the tanks and reservoirs located where there is

no danger of puncture from a collapsing landing gear?

4. Are the tanks and reservoirs located where transmissions, engines, and similar massive components

will not crush them during a crash?

5. Are the tanks and reservoirs relatively safe from

penetrative damage by structural stringers and

stiffeners?

6. Can the oil tanks displace in the airframe structure

and still not leak around the filler area, the fluid

line enitry and exit, the quantity indicator, and the

tank-to-structure attachment points?

7. Are the hydraulic reservoirs constructed and mounted

to withstand 30-G forces applied in any direction?

6.9.2.2 Oil and Hydraulic Lines

8. Are all oil and hydraulic lines made from flexible

hose with a steel-braided outer sheath wherever

possible?

9. Do all hose assemblies meet the strength requirements listed in Table 17, Section 6.2.3.1?

10. Can all hoses elongate 20 percent without the hose

assemblies spilling fluid?

11. Is coiled metal tubing used in areas where flexible

hose cannot be used, but large structural deformations are expected?

12. Has the number of fluid lines in the engine compartment been held to a minimum?

161
Yes No N/A

13. Are fluid lines routed along heavier structural

members wherever possible?

14. Are fluid lines routed as far as possible from

occupiable areas and electrical compartments?

15. Are fluid lines routed as far as possible from

all electrical equipment and wires?

16. Are fluid lines routed away from areas where large

structural damage is likely during a crash?

17. Are fluid lines routed away from the exhaust

system and high-temperature heating ducts?

18. Are the fluid system lines designed with as few

fittings as possible? *

19. Are the fluid system lines designed so that continuous hoses are run through bulkheads rather

than attached to the bulkheads with fittings? -

20. Are self-sealing breakaway valves used wherever a

fluid line goes through a tirewali or a bulkhead or

is attached to the bulkhead?

21. Are self-sealing breakaway valves used at all points

in the fluid lines where aircraft structural deformation could lead to line failure?

22. Are fluid line supports frangible to ensure release

of the line during crash impact?

23. Are uncut lines running through bulkheads stabilized

6.9.2.3 0tl and Hydraulic System Components

24. Are all oil and hydraulic system components located

as far as possible from anticipated impact areas,

occupiable areas, and electrical compartments?

25. Are the components located in the engine compartment

restricted to those absolutely necessary for engine

operation?

26. Can the construction and mounting of all system

components withstand 30-G forces applied in any

direction without leakage?

162
Yes No N/A

6.9.2.4 01 Coolers

27. Is the oil cooler located outside of the engine

compartment?

28. Is the oil cooler located as far as possible from

anticipated impact areas, occupiable areas, and

other potentially injurious components?

29. Can the oil cooler and connecting lines experience

considerable deformation without leaking?

30. Can the oil cooler mounting withstand 30-G forces

applied in any direction?

6.9.3 Ignition Source Control Checklist

6.9.3.1 Electrical Systems

1. Are wires routed as high up in the structure as

possible?

2. Are wires routed away from areas of anticipated

structural damage, i.e., landing gear failure,

nose crush-in, etc.?

3. Are wires routed above or away from flammable

fluid lines?

4. Are all wires routed through the structure so

that extensive structural collapse or displacement

can take place without breaking wiring?

5. Are wire bundles supported at frequent intervals by

Trarigible attach.e.ts to ..h. a.ir.r Jt structui-e?

6. Are wires shielded by felt or similar protective

covers in areas where crushing is likely?

7. Are wires to electrically operated boost pumps

20 to 30 percent longer than necessary?

8. Is all electrical wiring going through the fuel

tank compartments shrouded?

9. Is wiring in the fuel tank compartment routed as

high as possible in the compartment?

10. Are electrical wires in the fuel tank compartment

20 to 30 percent longer than necessary?

163
Yes No B/A

11. Are batteries, generators, and inverters located

in areas relatively free from structural collapse?

12. Are batteries, generators, and inverters located

as far as possible from flammable fluids?

13. Are batteries and generators (unless engine mounted)

housed in compartments built into the airframe?

14. Are battery, inverter, and generator mountings

capable of withstanding a 30-G force applied in

any direction?

15. Are the wires connecting the generator, battery,

and inverter into the system located in relatively

crush-free areas?

16. Are light bulbs and attaching wires on lower airframe surfaces designed to readily displace, rather

than remain stationary and be broken?

17. Are all electrical compartments lined with a

tough, nonconductive paneling?

6.9.3.2 Shielding

18. Are fuel tanks isolated from the occupants by a

minimum of two spillage barriers?

19. Are firewalls designed to withstand all survivable

crash impacts without losing their structural

integrity or sealing ability?

20. Are drainage holes located in all flammable fluid

tank compartments?

21. Is the hot metal of the engine shielded from flammable fluid spillages?

6.9.4 Interior Materials Selection Checklist

1. Do all interior materials meet the flammability

requirements specified in Federal Air

Regulation (FAR) 25.853?

2. Do all interior materials produce the lowest

possible amount of smoke and toxic gases (see

Section 6.5.3 for appropriate screening tests)?

164
Yes No N/A

6.9.5 Ditchlng Provisions Checklist

1. Are emergency exits larger and more numerous than

normally required to meet minimum standards?

2. Are additional escape exits provided in the

overhead, deck, and tail sections?

3. Have explosively created exit systems been

considered?

4. Are emergency exits lighted with high intensity

lights with a minimum brightness of 120 fL?

5. Even though escape lights meet the minimum requirement, is the brightness level of escape lighting

the highest permitted by other design conditions

(up to 200 fL)?

6. Has more than one aircraft flotation method been

provided?

7. Does the flotation bag system have a high

reliability?

8. Are tiedown or stowage facilities provided for

life rafts and other ditching equipment?

9. Are equipment restraint devices and supporting

structures designed to restrain the equipment to

loads of 50 G downward, 10 G upward, 35 G forward,

15 G aftward, and 25 G sideward?

10. Is all survival equipment readily available and

easily released after ditching?

11. Can life rafts be removed and deployed outside

the aircraft within 30 sec?

6.9.6 Emergency Escape Design Checklist

6.9.6.1 Emergency Exits

1. Are the numbers, sizes, and locations of the exits

such that a full load of troops and crew can evacuate in 30 sec when the aircraft is on its side?

2. Are all escape exits a minimum of 22 in. in diameter, or 22 in. square with 6-in. radius corners?

165
Yes No N/A

3. Can all emergency exits be completely opened wthin

5 sec after the person initiating the action first

places his hand on the release handle?

4. Does each crew mewber h, . access to at least one

eme;',.ncy exit regardless of aircraft attitude?

5. Are a minimum of two exits, one on ea:h side of the

fuselage, provided in troop/passenger compartnents?

6. Is at least one exit provided for every 10 perscns

e)xpected to occupy troopipassenger compartments?

7. Are emercency cxit locations equally divided on

each sdQd of the aircraft?

8. If the w'idLh of the fuselage is 5 ft or more, are

additiunal exits provide@ in the overhead, botiom,

fo,'e or aft section; of the airc,'aft?

9. Are all exit releasc mechani.•r.is of the sinale motion

typO:?

10. Is the numiber of different types of exit release

ha:,dles Ould to a minimum?

11. Can all exits be opened fromi both the inside and

outside uf the aircraft?

12. Can the exits be opened even if the fuselage

evidences tconsiderable distortion?

13. Can th,, exits be easily operated wheo' the aircraft

i, (n its side?

14. Wil I romuved ur upecncd exit covers inherently be l•;-.,i .... 'su as to) r,, blocl• the" evit roprninug. Ui i. ý--j a,, U, -u I, b 0,e , t ' ...... -,,- -

nur intiurfer.-i with, uccupatit egress?

)5. I , tOIh exit opU lirii olupraLion designed to inh.!r-ettly rVQ!st JaI1,,,l43 by lU!o(. 0I~j(:-LtS?

16. Can an exit bee vpened easily when the op).r atur is

L!I'iv iu'jhid or cruw(ILd by UIthV occup)ants?r -

P DV)urV i t!) Im-rg'rhCy eVdUaut i ur', du ad I l pas seniyrrs lave

eCent tii-,ly the :am(-: dlstance t( move during egress?

lo(,
Yes No N/A

18. Are aisles between seat rows wide enough to allow

unobstructed movement of occupants (at least 17 in.

minimum)?

19. If occupants mu:t pass through seat rows to reach

the exits, can they move to the exits at a rate

tý&t permits one person to exit every 1.5 sec

or less?

6.9.6.2 ExDlqsive Exit SSstems

20. Arc arming and firing accomplished in two separate

and deliberate actions?

'21. Is the arming function under the control of the

fliqht crew?

22. Will the Laf./arm mechanism r.main in its preselected poaition regardless of system failure

or environinuntal or crasi- inputs?

23. is the firing mcclanism independent of any external

eheroy ~,ource?

24. Can the exits hc opnried indepQl(:dently of each other?

2$. Are tOe explosiwe chargps u:..ed Lo cut the openings

hPld secut-ly in positiour aqJ3inSt the aircrafL

structure?

26. A,'e entrgy-abborbing .acku'y rwaterials placed behind

the explosi- chargts?

27. Car( thc s.vs•or,- fun:c.ion in arriLiont air temperature

up to 40C , Yet ;ot fun(c tion during 30- to

60. sec exp:sures t1, po'.tirash f-''es?

/'. t,,he a.,iurt ay.i,, dur;,t"on of aoy exposed fl ames

fron explos ivt actuati 'n •i iniirizl?

29. DUes the Interior viemergency 1 ightirig provide suf fi -

cic,1l. ill ,•,InatiU, to permit ULcupants to lu,.atu

ene,'y.•;cy xxlts, survival equipment, arid escape

path.?

3, Is there an average illurn iation In .lcear 4tir of

0.05 Qc or greater rnca:dred 20 in. aouvu thel floor

alony pissagyways leadinU to exits?

167 t7
Yes No A

31. Are supplementary lighting units located in the

lower part of the cabin at or near each emergency

exit?

32. Do all internally illuminated exit signs have a

minimum brightness of at least 25 fL.?

33. For noncombat missions, is exterior emergency

lighting provided to illuminate the ground near

each exit and the areas where escape and survival

equipment will be deployed?

34. Is the exterior light intensity on the ground ax

least 0.02 fc?

35. Can the lighting system withstand the crash conditions listed in Section 6.7.2.4, and still function? *

36. Is emergency lighting power independent of aircraft

power systems?

31. Can the emergency lighting system be actuated both

automatically and manually?

6.9.6.4 Emergency Exit Markings.

38. Are emergency exits clearly marked both inside and

outside the aircraft?

39. Are instructions for releasing the exits clearly

marked beside the exit release mechanisms?

40. Do all exit mar!.ins meet the requirunments of the

Department of Army Technical Bulletin 746-93-2?

1. Can the (:rash locator beacor, be activated both

automatically arid manually?

2. Is an inertia sens,)r used to automatically activate thi: br-a:,,r*'

3. IDo l! lo(vjituiji.i•,I ,;.ertla !,h.iors in fixed winrj

,l Y r, i' rwt ,e,. ".he •r t u,,Li£ lp ,loit', stated in

4. '.• Ol' )o~ily 1 ttl'lt,, l .!.(J vlJt ic.,l i,'',Uti '). 11rsors

1'VUV ov dr ej• i ii(aft? n v ii 

IIfi
5. Do the inertia sensors in rotary-wing aircraft meet

the actuation limits stated in Section 6.8.2?

6. Is the inertia sensor mounted solidly to rigid

structure located in an area that will experience

crash forces representative of those in the

occupant compartment?

7. Are the transmitter ad antenna loc.ted in areas

that are least subject to impact damage?

8. Can the transmitter and antenna withstand the

crash forces listed in Section 6.7.2.4?

9. Does the crash locator beacon have its own

independent power supply'?

10. Is al", electrical wiring between system components protected from impact damage?

I BI

Ui

Ib
REF EREiNC ES

i. Military Stlandard MIL-STL:-1290A(AV), LIGHT FIXED- AND ROTARY-WING AIRCkAFT CRASH1 RESISTANCE, Departme~nt of Defense, W;ýsh;,.-ýton, DC, 20301,

26 ýqitembex M~8.

2. ENGINECR~IG ANALYSIS OF Cl-MSH INJUR~Y IN AI~r4Y OFI-58 AIRCRAFT, USASC Technical Rep,.irt 79-1, U, S,~ Army Safety Center, Fort. Rucker, AL, January

19,419

3. E.NGINEERING~ ANIALYSIS GF CRAS'91 !NJUfkY IN ARM~Y (Ai-47 AIR'.RAFT, USAAAVS TechP'cal RcportA 784 U. S. Ariiy Agency for Aviation Safety, Fort Pucker,

AL, June 1978.

4. ENGINEERING ANALYSIS OF CRA~SH INJUR': IN~ ARMY AN-1 AIRCRAFT, USAAAVS Techn4 ' a] Report 783-3, U. S. Army Agency for Aviationi Safety, Fort Rucker,

A,-. Mar'ch '1b71S.

5. Carne.fl, f). L., C[RASHtIORTHME4SS DESIGN FEATURES FU' ADVANCED U11ILITY HELICO0PTERS, -:n A.ir.-raft Cr,30-wort-Vhness, K. Siiczal ski et al]., eds., University rrs! L) 'irginia, Charloltesvil k, VA, 1'7,pp. 51-64.

6l .',r~ jrbxidge, M. ., R~eilly, M. J. , .ý,md Cori s zlvos, J. F. , CRASHWORTH INESS

OiF THE ý.0JEMC VEITOL UTTAS, in .rcat -ah rhns, K. Saczal ski et

al., ell~s. duvrst 'r.sf iryinia, C1ar1"ittes;ville, VA, 1975,

7. P (:1, M. J. , IH4L Sf IGAM iOFJo ADVANCED HEl .ICOP~i fR STRUCTURAL DESIGNS,

Vol mrc 1, iAUVAq'1LD TN(7U<L 1 M)OUNE t DES I.Vi CU.ICFLPT STUE.Y, Sikorsky

Aircr-itft, bi 'i sior, 3F (ihited hýchn'i ogy Corpo-, at-Ion, USAAMRDL Technical

I~~r Uu1sli *~9A W-irtor~l,-, U). S. f'rrry Air Mobility Presearch and

f~~'t f~1121t;. aI(iratiry , _ort l~r'-t s, V, I*J 1976, Al

1;. hoI1ttD ý,I and Sw~tton1, S . ,AUVANCUtJ 1PC R.~ESIGN Oi1Ca 

i~~ivr 1~Tsr 1 1..1 6(B i fM ýr tol Cornpdvy ; USM¶ikDI. 1(ý b i cal I, .

~' ~.5Eh.,L~tr; O~t ~rU. S. Aý"n.>ý Ar M1rjbi1 ty R''rhi,';d( f'ovelop- mci1 cry,! '' . VA, Karrh !?~'/(, AQ P46:.2.

9. Ick, J. i. , Ull ANP .'IS) Or, I I Ecf fCY .l.ArC] DL. i COSI S 10'r 1 M: AV'~.K

Si-U1 :m r U.T L S. r,-,y Aiti. fur Aviat ion Safely. Fcr t Rucb rr, AL.,

i:0. 14Jr''L,3 J, ae' 'l, f.., lI1S VI iIFiS V I AUtSl N11.IiANY(:irA

;i.' HFAy, J. I.., Jr 'TIWRliiF{ vk C : A SiIVt)k (A I.'IMY VUOI/'>l

WI'G ~h(.A IAs JIt 1 ~J~IN FPIR10 JAtkARY 19*i/0 1HP.Ur'd l;I. Irwlii 191!

-)11, :r f) o i:nt ud I t h c Ali vcr ? f t Cr sf r t I 'i~2;~ Ilp I u11i, U1-:Ill v .iif y

I 1 CA ,15 I , 011, 0" obt r 
REFERENCES (CONTD)

12. H~cks, J. E., ECONOMIC BENEFIIS OF UTILITY AIRCRAFT CRASHWORTHINESS,

USAAAVS Technical Report 76-2, U. S. Army Agency for Aviation Safety,

Fort Rucker, AL, July 1976.

13. TEE ECONOMIC BLNEFITS OF CRASHWORTHINESS AND FLIGHT SAFETY DESIGN FEATURES IN ATTACK HELICOPTERS, USAAAVS lechnical Report 77-2, U. S. Army

Agency for Aviation Safety, Fort Rucker, AL, June 1977.

14. Cronkhite, J. D., DESIGN OF AIRFRAME STRUCTURES FOR CRASH IMPACT, Bell

Helicopter Textron, Inc., presented t the American Helicopter Society

National Specialist's Meeting on Crashworthy Design of Rotorcraft at

Georgia Institute of Technology, Atlanta, GA, April 7-9, 1986.

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NON-EJECTION, AIRCREW, GENERAL SPECIFICATION FOR, Department of Defense,

Washington, DC, 31 January 1986.

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20301, 19 September 1981.

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rA U-,O,,lYuv FUEL SYSTEM IN A CRASH ENVIRONIMENT, DY..m. Science, Divi

sion of Marshall Industries, USAtMRBL lechnical Report 71-1.7, Eustis

Directorate, U.S. Army Air Mobility Research and Development Laboratory,

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19. Zimunerrnann, Richard E., RETROIII ENERGY-ABSORBID'S CREWSEAT FOR THE S11-3

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20. Share, S., J, and Carnull, B. L.., lifl DESIGN AND QUJALIFICA OI [i0:q SI ING

01 AN ENERGY ABSORbING SLAl I'OR 111[ IIAV''S 14 53A/D HILLICO1IEL., Simule

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